THE  LIBRARY 

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Pacific  Aeronautical  Library 


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AEROPLANE 

CONSTRUCTION  AND 

OPERATION 

Including  Notes  on  Aeroplane  Design 

and  Aerodynamic  Calculation, 

Materials,  Etc. 


A    Comprehensive    Illustrated    Manual    of    Instruction    for 

Aeroplane  Constructors,  Aviators,  Aero-jVIechanics, 

Flight  Officers  and  Students.     Adapted 

Either  for  Schools  or  Home  Study. 


BY 

JOHN  B.  RATHBUN 

AERONAUTICAL  ENGINEER 

Consulting  Aeronautical  Engineer,  Chicago  Aero  Works:  Chief  Engineer 

Automotive     Engineering    Company.       Formerly     Instructor    in 

Aviation  and  Machine  Design,  Chicago  Technical  College. 


DISCARDED  B\^—       P.A.L. 


CHICAGO 

STANIQNan^VANWETS. 

PUBLISHERS 

1918 


Z-Ol  \A/ 


Copyrighted,  1918 
By  STANTON  AND  VAN  VLIET  CO. 


DISCARDED  BY  P.A.L. 


INTRODUCTION 

Many  aeronautical  books  of  a  purely  descriptive  nature 
have  been  written  for  the  average  man,  but  as  a  rule  they 
contain  little  of  interest  for  the  more  serious  student  of 
the  subject.  Other  books  of  a  highly  technical  and  mathe- 
matical class  have  also  been  published,  but  their  contents 
are  all  but  unintelligible  to  anyone  but  a  trained  engineer. 
It  is  the  purpose  of  the  author  to  compromise  between 
these  two  extremes,  and  give  only  that  part  of  the  theory 
and  description  that  will  be  of  practical  use  for  the  builder 
and  flyer.  The  scope  of  the  subjects  covered  in  this 
volume  has  been  suggested  by  the  questions  asked  by 
students  and  clients,  and  is  the  result  of  many  years'  cor- 
respondence with  beginner  aviators  and  amateur  aero- 
plane builders. 

I  have  endeavored  to  explain  the  principles  of  the  aero- 
plane in  simple,  concise  language,  starting  with  the  most 
elementary  ideas  of  flight  and  finishing  with  the  com- 
plete calculations  for  the  surfaces,  power,  weight,  etc. 
When  mathematical  operations  are  necessary  they  are 
simple  in  form,  and  are  accompanied  by  practical  prob- 
lems worked  out  numerically,  so  that  a  man  with  even  the 
most  elementary  mathematical  knowledge  will  have  no 
difficulty  in  applying  the  principle  to  his  own  work.  In 
cases  where  the  calculations  would  necessarily  be  compli- 
cated, I  have  substituted  tables  of  dimensions  for  the 
mathematical  operations,  these  dimensions  being  taken 
from  a  number  of  representative  machines. 

While  flying  cannot  be  taught  by  books,  and  is  only  the 
result  of  actual  experience,  the  chapter  devoted  to  the  use 
of  controls  under  different  flight  conditions  will  be  of  great 
benefit  to  the  prospective  aviator.    The  portion  of  the  book 


INTRODUCTION 

devoted  to  operation  will  be  of  use  in  flying  schools  and 
training  camps  since  both  training  methods  and  control 
manipulation  are  covered  in  detail.  In  addition  I  have 
presented  considerable  data  on  the  requirements  of  the 
modern  aeronautical  motor. 

So  many  new  firms  are  now  entering  the  aeroplane 
industry  that  there  is  an  ever  increasing  demand  for 
trained  mechanics,  designers  and  flyers,  and  many  tech- 
nical men  now  working  along  other  lines  are  taking  a  keen 
interest  in  aeronautical  engineering.  If  the  contents  of 
this  book  will  serve  toinspirethe  technical  readerto  deeper 
interest  and  practical  research  in  the  fascinating  subject 
of  aeronautics,  the  author  will  be  more  than  satisfied  with 
the  result  of  his  labor.  The  aeroplane  is  rapidly  assuming 
a  great  commercial  importance,  and  there  is  no  doubt  but 
what  it  will  develop  into  an  industry  rivaling  that  of  the 
automobile. 

To  keep  fully  abreast  of  the  times  in  aeronautic  develop- 
ment, one  should  be  a  constant  reader  of  the  excellent 
aeronautical  magazines.  Too  much  praise  cannot  be  given 
to  the  aeronautical  press  in  its  effort  to  maintain  an 
interest  in  this  subject,  and  as  with  all  pioneering  move- 
ments, these  magazines  have  met  with  many  discourage- 
ments and  financial  setbacks  in  the  earlier  days  of  flying. 
To  the  American  magazines,  "Aerial  Age"  and  "Flying" 
(New  York),  the  author  owes  a  debt  of  gratitude  for  the 
use  of  several  of  the  cuts  appearing  in  this  book.  The 
English  magazines,  "Flight,"  "Aeronautics"  and  the 
"Aeroplane,"  have  been  similarly  drawn  on.  "Aviation 
and  Aeronautical  Engineering"  (New  York)  has  sug- 
gested the  arrangement  of  several  of  the  tables  included 
herein.  All  of  these  papers  are  of  the  greatest  interest 
and  importance  to  the  engineer,  aviator  and  aero-me- 
chanic. 

JOHN  B.  RATHBUN. 


TABLE  OF  CONTENTS 


CHAPTER  I.  Aeroplane  Principles.  The  Aeroplane,  Or- 
thopter,  Helicopter,  Ornithopter  Dynamic  Flight,  Mono- 
plane, Biplane,  Triplane,  Operation  Principles 7 

CHAPTER  II.  Military  Aeroplane  Types.  Service  Classi- 
fication, Chasers,  Bombers,  Reconnaissance  Types,  Bat- 
tle Planer,  Trainers,  Tables  of  General  Dimensions. 
Power,  Weight,  Etc.,  Comparative  Speeds,  Climb,  and 
Capacity    37 

CHAPTER  III.  Elementary  Aerodynamics.  Properties  of 
Air,  Altitude.  Density.  Pressure,  Effect  of  Altitude  on 
Power  and  Lift,  Air  Tables.  Pressure  on  Flat  Plates, 
Inclined  Flat  Planes,  Normal  Plates,  Streamline  Bodies, 
Turbulence,  Skin  Friction,  Aspect  Ratio,  Lift  Calcula- 
tions, Similitude  71 

CHAPTER  IV.  Experimental  Laboratories.  Aerodynamic 
Test  Methods,  Wind  Tunnels.  Whirling  Arm  Method, 
Full  Size  Tests,  Eiffel,  N.  P.  L.,  and  AI.  I.  T.  Laborator.     93 

CHAPTER  V.  Aerofoils.  Lift  by  Curved  Plates,  Camber, 
Air  Flow  About  Aerofoil,  Lift  and  Drag  Components, 
Chordal  Line,  Incidence,  Critical  Angle.  Thick  Wings, 
C.  P.  Movement,  Pressure  Distribution,  Aerofoil  Types, 
Forms  of  Charts,  Lift-Drag  Ratio,  Effects  of  Camber, 
Eiffel's    Results    100 

CHAPTER  VI.  Practical  Wing  Sections.  Practical  Re- 
quirements of  Wing  Sections.  Camber  Proportions, 
L/D  Ratios,  Loading,  Structural  Requirements,  Tables 
and  Charts  of  Wing  Performance  for  RAF-3-6,  Eiffel- 
32-36-37,  USA-1-2-3-4-5-6,  and  Curtiss  Sections.  Wing 
Correction  Factors  123 

CHAPTER  VII.  Biplanes  and  Triplanes.  Superposed 
Planes,  Interference,  Biplane  and  Triplane  Reduction 
Factors,  Gap,  Gap-Chord  Ratio,  Stagger,  Decalage, 
Form  of  Opposing  Wing  Surfaces,  Triplane  Reduction 
Factors,  Stability 165 


CONTENTS 

PAGE 

CHAPTER  VIII.  Effects  of  Plan  Form.  Wing  Outlines 
and  Arrangement,  Sweep  Back,  Rake,  Negative  Tips, 
Tandem  Wings,  Stability  Due  to  Sweep  Back 181 

CHAPTER  IX.  Wing  Construction.  Arrangement  of  Wing 
Members,  Spars,  Ribs,  Fabric,  Drag  Bracing,  Edging, 
Types  of  Wings  and  Wing  Members  Illustrated,  Spar 
Loading,   Effect  of  C.  P.,   Wing  Fittings,   Spar  Types, 

Sub  Ribs 191 

• 

CHAPTER  X.  Wing  Construction  Details.  Rib  Construc- 
tion, Stiffeners,  Battens,  Rib  Strength,  Tests,  Forming 
Ribs,  Spars  and  Construction,  Box  Type,  "I"  Beam,  Com- 
posite, Tube,  Strut  Connections,  Protection,  Spar  Lo- 
cation      210 


CHAPTER  XI.  Fuselage  Construction.  Types  of  Fuselage, 
Wired  Truss,  Monocoque,  \'eneer,  Steel,  Resistance,  Lo- 
cation of  Passengers,  Motor.  Etc.,  Center  Line  of  Re- 
sistance, Motor  Arrangement  Illustrated,  Instruments, 
Dimensions    227 


CHAPTER  XII.     Details  of  Fuselage  Construction.    Details 

of  Wiring,  Wing  Connections,  Strut  Fittings,  Weights, 
Longeron  Dimensions,  Steel  Frames 253 

CHAPTER  XIII.  Chassis  Construction.  "V"  Type,  Skids, 
Bleriot,  Miscellaneous  Types,  Wheels,  Hubs,  Tires, 
Shock  Absorbers,  Location  of  Wheels,  Tail  Skids,  Land- 
ing and  Running  Shock,  Illustration  of  Details 287 

CHAPTER  XIV.  Estimation  of  Weight.  Useful  Load, 
Dead  Load,  Load  Distribution,  Weights  by  Percentage, 
Itemized  Weights,  Tables,  Loading  and  Speed,  Influence 
on  Climb,  Power  Plant  Weight,  Calculations 299 

CHAPTER  XV.  Balance  and  Stability.  Static  and  Dynamic 
Stability,  Centers  of  Gravity,  Thrust,  Lift  and  Area, 
C.  P.  Movement,  Damping,  Stabilizing  Surfaces,  Control 
Surfaces,  Roll,  Pitch,  Yaw,  Three  Axes,  Surface  Calcu- 
lations, Ailerons,  Control  Systems,  Inherent  Stability, 
Calculating  C.  G 311 

CHAPTER  XVI.  Head  Resistance.  Structural  or  Parasitic 
Resistance,  Total  Head  Resistance  Calculations,  Re- 
sistance of  Wires,  Struts,  Wheels,  Radiator,  Etc.,  Re- 
sistance Distribution  by  Percentage,  Gliding  Angle,  Body 
Resistance    337 


CONTENTS 

PAGE 

CHAPTER  XVII.  Power  Calculations.  Variations  of 
Power  with  Speed  and  Resistance,  Normal  Speed, 
Limiting  Speeds,  Power  Charts,  Propeller  Efficiency, 
Power  and  Gliding  Angle,  Power  for  Climb,  Power 
Calculations 355 

CHAPTER  XVIII.  Propellers.  Slip  Stream  Reaction, 
Pitch  Speed,  Slip,  Thrust,  Diameter,  Blade  Outline, 
Blade  Section,  Uniform  Pitch,  True  Screw  Type,  Diam- 
eter and  Wing  Span,  Horsepower  Absorbed,  Construc- 
tion Notes   366 

CHAPTER  XIX.  Operation  and  Training  Methods.  Ele- 
ments of  Flying,  Training  Methods.  Dual  Control,  Pen- 
guins, Self  Training,  Rolling,  The  First  Solo,  Handling 
the  Controls,  Instructions  for  Self  Training,  Rudder  Con- 
trol, Elevator  Control,  Ailerons,  Adjustments  for  First 
Flight  Alone.  The  First  "Straight"  Landing,  Turns  and 
Banking.  Side  Slip,  Stalling,  \'ol  Plane.  Stunt  Flying, 
Spiral,  Tail  Spin,  Skidding,  Immelman  Turn,  Operation 
in  General ^7S 

CHAPTER  XX.  Aeronautic  Motors.  Desirable  Character- 
istics of  Aeronautic  Motors,  Weight,  Fuel  Consumption, 
Compression,  Duty.  Four  Stroke  and  Two  Stroke  Cycle 
Types,  Rotating  Cylinders.  Power  Estimates,  Influence 
of  Altitude,  Tables  of  Capacity,  Weight,  and  Dimensions 
of  Typical  Motors 387 

CHAPTER  XXI.  Glossary  of  Aeronautic  Words  and 
Phrases.  An  alphabetically  arranged  list  of  the  most 
commonly  used  aeronautic  words  and  phrases  with  their 
French  Equivalents.  It  is  arranged  so  that  it  can  be  used 
as  an  English-French  or  French-English  glossary,  and 
contains  many  words  that  can  not  be  found  elsewhere. 
The  new  Advisory  Committee  words  are  included 403 


CONTENTS 


AERONAUTICAL  MAGAZINES 

The  following  list  of  American  and  English  aeronautic  publications  will  be 
of  interest  to  those  who  wish  to  keep  in  touch  with  the  latest  developments  in 
aeronautics: 

AVIATION  AND  AERONAUTICAL  ENGINEERING  (two  issues  per  month). 
.     A  technical   magazine  published  by  The  Gardner-Moflfat   Co.,  Inc.,    120  W. 
32d  St.,  New  York. 

AERIAL  AGE  (weekly).     Popular  and  technical.     The  Aerial  Age   Co.,  Foster 
Bldg.,  Madison  Ave.  and  40th  St.,  New  York. 

AIR  SERVICE  MAGAZINE  (weekly).      Military  and  popular  subjects.     Gard- 
ner-Moffat  Co.,  Inc.,  120  W.  32d  St.,  New  York. 

FLYING    (monthly).      Popular    and    military    subjects.      Published    by    Flying 
Association,  Inc.,  280  Madison  Ave.,  New  York. 

AIR  TRAVEL    (weekly).      Popular   subjects.      Published   by    Air   Tiavel,    New- 
York. 


ENGLISH  MAGAZINES. 

FLIGHT  AND  THE  AIRCRAFT  ENGINEER  (weekly).  Technical  and  popu- 
lar. Published  by  Flight  and  Aircraft  Engineer,  36  Great  Queen  St., 
Kingsway,  W.C.2,  London,  England. 

AERONAUTICS  (weekly).  Technical  and  industrial.  Published  by  Aero- 
nautics, 6-8  Bouverie  St.,  London,  E.G. 4,  or  may  be  had  from  1790  Broad- 
way, New  York. 

THE  AEROPLANE  (weekly).  Technical  and  popular.  Published  by  "The 
Aeroplane,"  166  Piccadilly,  London,  W.l. 


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CHAPTER  I. 
PRINCIPLES  OF  THE  AEROPLANE. 

Mechanical  Flight.  Although  the  elementary  principles 
of  mechanical  flight  are  not  of  recent  origin,  the  practical 
development  of  the  flying  machine  is  confined  almost  en- 
tirely to  the  present  century.  Gravity  propelled  gliders 
and  small  models  have  been  flown  with  success  from  a 
comparatively  early  date,  but  the  first  actual  sustained 
flight  with  a  power  driven  machine  was  performed  by  the 
A\>ight  Brothers  in  1903.  There  was  no  single  element 
on  this  first  successful  machine  that  had  not  been  proposed 
many  years  before  by  Langley,  Chanute,  Montgomery, 
Henson,  Mouillard,  and  others,  but  this  first  flight  must 
be  attributed  principally  to  the  fact  that  the  Wrights 
started  carefully  and  painstakingly  to  learn  how  to  operate 
(By  practicing  with  gliders)  before  starting  on  the  first 
power  machine.  If  Langley  had  studied  the  operation  of 
his  machine  as  carefully  as  he  did  its  theory  and  design, 
he  would  have  been  flying  long  before  the  Wrights  as  his 
original  machine  was  afterwards  successfully  flown  by 
Curtiss. 

\\'hen  once  actual  flight  was  achieved,  and  the  success 
of  the  Wright  Brothers  became  generally  known,  the  de- 
velopment proceeded  with  leaps  and  bounds.  All  the  re- 
sources of  science  and  engineering  skill  were  at  once  ap- 
plied to  the  new  device  until  our  present  scientific  knowl- 
edge of  the  aeroplane  compares  very  favorably  with  the 
older  engineering  sciences.     In  the  few  years  that  have 

7 


8  AEROPLANE  PRINCIPLES 

elapsed  since  the  first  flight,  the  aeroplane  holds  all 
records  for  speed,  endurance,  and  radius  of  action. 

A  great  deal  of  the  success  so  rapidly  acquired  can  be 
credited  to  the  automobile  and  motorcycle  industries, 
since  it  was  the  development  of  the  light  internal  combus- 
tion motors  used  on  these  machines  that  paved  the  way 
for  the  still  lighter  aeronautic  motor.  Again,  the  auto- 
mobile industry  was  responsible  for  the  light  and  powerful 
materials  of  construction,  such  as  alloy  steel,  aluminum 
alloys,  and  also  for  the  highly  important  constructional 
details,  such  as  ball  bearings,  pneumatic  tires,  carburetors, 
magnetos,  steel  tubing,  etc.  The  special  methods  de- 
veloped in  automobile  work  have  helped  to  make  the 
aeroplane  an  immediate  commercial  proposition. 

Types  of  Flying  Machines.  In  general,  flight  apparatus 
may  be  divided  into  two  classes,  (1)  The  Lighter  Than 
Air  Type,  such  as  the  balloon  and  dirigible,  and  (2)  The 
Heavier  Than  Air  Machine,  represented  by  the  aeroplane, 
helicoptor  and  ornithopter.  The  lighter  than  air  machine 
is  supported  in  flight  by  ''bouyancy"  in  much  the  same 
manner  that  a  piece  of  wood  floats  in  water.  When  a 
balloon  or  dirigible,  because  of  its  large  volume,  displaces 
a  volume  of  air  equal  to  its  own  weight,  the  device  will 
float.  When  the  weight  of  air  displaced  exceeds  the 
weight  of  the  balloon  or  dirigible,  it  will  continue  to  rise 
until  it  reaches  an  altitude  where  the  diminished  air  density 
again  results  in  an  equality  between  the  weight  of  the 
device  and  the  air  displaced.  At  this  point  it  rests,  or  is  in 
equilibrium.  The  flotation  of  such  a  device  is  entirely  due 
to  static  forces  and  hence  (1)  is  often  called  an  "aerostat." 

The  sustenation  of  a  Heavier  Than  Air  Machine  is  due 
to  an  entirely  different  application  of  forces.  Forces  in 
motion  (Dynamic  Forces)  are  essential  to  the  support  of  a 
heavier  than  air  machine,  and  it  is  the  resultant  of  these 
forces  that  performs  the  actual  lifting  operation,  this  re- 
sultant corresponding  to  the  buoyant  force  of  the  aerostat. 


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10  AEROPLANE  PRINCIPLES 

''Dynamic"  flight  is  obtained  by  an  apparatus  in  which 
an  arrangement  of  surfaces  are  moved  in  such  a  way  as 
to  cause  an  upward  component  of  the  forces  generated  by 
the  impact  of  the  air  on  the  surfaces.  The  surfaces  drive 
the  air  down  and  when  the  force  necessary  for  the  con- 
tinuous downward  deflection  of  air  becomes  equal  to  the 
weight  of  the  machine  it  is  sustained  in  flight.  Dynamic 
flight  therefore  depends  on  the  continuous  downward  de- 
flection of  masses  of  air,  and  when  this  motion  ceases, 
sustentation  also  ceases. 

An  aeroplane  is  provided  with  a  deflecting  surface  that 
is  fixed  rigidly  in  regard  to  the  body  of  the  machine,  and 
the  motion  necessary  for  its  support  is  provided  by  driving 
the  machine  forward,  the  forward  motion  being  produced 
by  the  horizontal  pull  of  air  screws  or  propellers.  It  is  at 
once  evident  that  the  forward  horizontal  motion  of  the 
aeroplane  must  be  maintained  for  its  support,  for  the 
surfaces  are  fixed  and  there  is  no  other  possible  way  of 
producing  a  relative  motion  between  the  wings  and  the  air. 

To  overcome  the  objection  of  forward  motion,  several 
other  machines  have  been  proposed  in  which  the  surfaces 
are  moved  in  relation  to  the  body,  as  well  as  the  air, 
thus  making  it  possible  for  the  device  to  stand  stationary 
while  the  revolving  or  reciprocating  surfaces  still  con- 
tinue in  motion  in  regard  to  the  air.  One  type  of  the 
moving  surface  machine,  the  "Helicopter,"  is  provided 
with  revolving  surfaces  arranged  in  the  form  of  vertical 
air  screws  or  propellers,  the  blades  of  the  propellers  being 
inclined  so  that  they  drive  down  a  continuous  stream  of 
air  and  produce  the  continuous  upward  reaction  that  sup- 
ports the  machine.  While  such  machines  have  succeeded 
in  raising  themselves  off  the  ground  they  are  not  yet 
practical  flying  devices.  The  "ornithopter"  or  "orthop- 
ter"  is  a  flapping  wing  machine  that  maintains  flight  after 
the  manner  of  the  bird  (Ornis).  Like  the  helicopter,  the 
ornithopter  has  not  yet  proved  successful. 


AEROPLANE  PRINXIPLES 


11 


Principles  of  the  Aeroplane.  In  its  elementary  prin- 
ciples, the  aeroplane  can  be  compared  with  the  kite, 
as  both  are  supported  by  the  impact  of  a  horizontal 
stream  of  air.  In  Diagram  1,  the  kite  surface  is  indicated 
by  X-X  with  the  relative  air  stream  W-W-W-W  moving 
from  left  to  the  right  as  indicated  by  the  arrow  heads. 
On  striking  the  surface,  the  air  stream  is  deflected  ver- 
tically, and  in  a  downward  direction,  as  shown  by  the 
streams  lines  R-R-R-R.  The  reaction  of  the  air  deflection 
produces  the  lift  shown  vertically  and  upwards  by  the 
arrow  L.     The  kite  surface  is  held  against  the  impact  ot 


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Fig  1.  romparison  Between  the  Kite  and  Aeroplane;  Fig.  2,  Showing  the  Lift 
and  Drag  Forces  Produced  bv  the  Air  Stream.  The  Propeller  (P),  Acts 
in  a  Manner  Similar  to  the  Kite  String  (S)  in  Producing  Relative  Mo- 
tion Between  the  Air  and  the   Lifting   Surfaces. 


the  air  stream  by  the  string  S  so  that  there  is  relative 
motion  between  the  air  and  the  kite,  and  so  that  the  sur- 
face will  not  be  carried  along  with  the  air  current  toward 
the  right.  If  the  kite  were  allowed  to  drift  with  the  wind 
there  could  be  no  relative  motion  between  the  surface  and 
the  air  stream,  hence  the  kite  would  fall  as  soon  as  it  at- 
tained the  velocity  of  the  wind.  The  horizontal  force 
exerted  by  the  wind  tending  to  carry  the  kite  toward  the 
right  is  indicated  by  the  arrow  D  and  is  known  as  the 
"drag"  or  "drift"  force.  There  are  thus  three  forces,  the 
lift  (L),  the  drag  (D),  and  the  resultant  of  the  two  forces 
indicated  by  the  string  (S).  The  forces  of  Hft  and  drag 
are  nearly  at  right  angles  to  one  another.    The  kite  tail  T 


12  AEROPLANE  PRINCIPLES 

is  simply  a  stabilizing  device  whose  purpose  is  to  maintain 
a  constant  angle  between  the  surface  and  the  wind  and 
it  performs  an  almost  negligible  amount  of  lift. 

A  few  more  words  in  regard  to  the  "relative  velocity" 
between  the  surface  and  wind.  In  the  figure,  the  kite  is 
assumed  as  being  stationary,  while  the  wind  moves  from 
left  to  right.  With  a  thirty  mile  per  hour  wind,  the  rela- 
tive air  velocity  in  regard  to  the  surface  would  be  30  M.  P. 
H.  If  the  air  particles  are  now  considered  stationary, 
and  if  the  kite  is  towed  toward  the  left  (opposite  to  figure) 
at  30  miles  per  hour,  the  relative  velocity  between  the 
surface  and  air  would  still  be  30  M.  P.  H.  In  other  words, 
the  kite  may  be  stationary,  or  may  move  in  regard  to  the 
earth,  but  its  lift  is  unaffected  as  long  as  the  relative  mo- 
tion between  the  surface  and  air  remains  constant.  The 
motion  between  an  aeroplane  and  the  earth  depends  upon 
the  dift'erence  of  the  aeroplane  and  wind  velocities.  For 
example,  a  aeroplane  with  a  relative  speed  of  60  miles  per 
hour,  flying  against  a  headwind  of  30  miles  per  hour, 
moves  60  —  30  ^  30  miles  per  hour  in  regard  to  the  earth. 
The  same  aeroplane  flying  with  the  above  wind  would 
have  a  velocity  of  60  +  30  =  90  miles  per  hour  past  a  fixed 
point  on  the  earth's  surface,  yet  in  both  cases,  the  relative 
velocity  of  the  surface  in  regard  to  the  air  would  be  the 
same. 

Fig.  2  is  a  diagram  of  an  aeroplane  that  shows  the  con- 
nection between  the  kite  and  aeroplane  principles.  In 
this  figure,  the  wing  surface  of  the  aeroplane,  X-X  cor- 
responds to  the  kite  surface  X-X.  The  relative  air  W-W- 
W-W  striking  the  wing  from  the  left  is  deflected  down 
along  the  arrows  R-R-R-R  and  results  in  an  equivalent 
lift  force  L,  and  a  drag  force  D  as  in  the  case  of  the  kite. 
The  resultant  force  required  to  maintain  the  relative  ve- 
locity between  the  air  and  wings  is  indicated  by  D^,  oppo- 
site and  equal  to  the  drag  force  D.  The  resultant  required 
for  overcoming  the  drag  is  provided  by  the  screw  pro- 


AEROPLANE  PRINCIPLES  13 

peller  P  instead  of  the  string  S  shown  in  Fig.  1.  The  pro- 
peller thrust  (D^)  is  parallel  to  the  air  stream  instead  of 
being  inclined  as  in  the  case  of  the  string,  but  the  total 
effect  is  the  same  since  both  are  "Resultants  of  the  lift 
and  drag."  To  sustain  the  aeroplane,  the  lift  (L)  must  be 
equal  and  opposite  to  the  weight  shown  by  M.  The  fact 
that  M  and  L  are  opposite  and  equal  makes  it  only 
necessary  for  the  propeller  to  overcome  the  horizontal 
drag,  and  hence  the  thrust  can  be  made  parallel  to  the  air 
flow — or  nearly  so.  The  aeroplane  is  provided  with  a 
small  tail  surface   (T)   that  corresponds  to  the  kite  tail 


Fig.  3.     Caudron    Monoplane.      Side    Elevation. 

(T).  It  maintains  the  lifting  surfaces  X-X  at  a  given 
angle  with  the  air  stream.  The  tail  may,  or  may  not  aid 
in  supporting  the  machine,  but  in  modern  machines  it  is 
common  to  employ  a  tail  surface  that  is  non-lifting  under 
ordinary  conditions  of  normal  flight.  The  body  (B)  con- 
tains the  pilot,  motive  power,  fuel,  and  such  useful  load 
as  it  may  be  necessary  to  carry. 

Fig.  3  shows  a  Caudron  monoplane  in  side  elevation. 
This  view  illustrates  the  application  of  the  principles 
shown  by  Fig.  2,  except  for  the  vertical  rudder  at  the  rear. 
The  latter  is  used  for  steering  in  a  horizontal  direction. 
Fig.  4  shows  the  construction  even  more  clearly  since  it 
is  a  perspective  view.  The  machine  is  a  Morane  "Parasol" 
monoplane  with  the  wing  placed  over  the  body.  This 
location  of  the  main  lifting  surface  is  for  the  purpose  of 


14 


AEROPLANE  PRINCIPLES 


improving  the  view  of  the  pilot  and  in  no  way  affects  the 
principles  just  described.  The  wires  shown  above  the 
wing  are  bracing  stays.  The  tail  is  hinged  near  the  rear 
so  that  the  angle  of  the  rear  portion  can  be  changed 
(Elevator  flaps),  and  permits  the  angle  of  the  main  wings 
to  be  altered  in  regard  to  the  air  stream,  thus  causing  the 
machine  to  ascend  or  descend.  The  tail  also  damps  out 
oscillations  or  vibrations  due  to  irregularities  in  the  air 
current.  The  wdieels  and  running  gear  (Chassis)  allow 
the  machine  to  be  run  over  the  ground  until  the  relative 
air  speed  is  sufficient  to  support  the  machine  in  flight. 


Fig.   4.      Morane   Umbrella   Type   Monoplane.  ,    Courtesy   of  "Flight." 

Multiplanes.  In  order  to  support  a  heavy  load,  and  at 
the  same  time  have  a  small  compact  machine,  it  is  neces- 
sary to  have  more  than  one  ''layer"  of  wing  surface.  It  is 
evident  that  the  wing  length  or  ''span"  can  be  made  much 
less  than  that  of  the  monoplane  surface  shown,  if  the  total 
area  could  be  divided  into  two  or  more  parts.  A  machine 
having  its  main  lifting  surface  divided  into  two  or  more 
separate  sections  is  known  as  a  "multiplane,"  this  term  be- 
coming "Biplane,"  "Triplane"  or  "Quadraplane,"  depend- 
ing on  whether  there  are  two,  three  or  four  independent 
lifting  surfaces.  There  is  almost  a  limitless  variety  of  ar- 
rangements possible,  but  the  most  common  arrangement 
by  far  is  that  of  the  biplane,  in  which  there  are  two  super- 


AEROPLANE  PRI^XIPLES  15 

posed  surfaces  as  shown  by  Fig.  5.  In  this  type,  the  two 
lifting  surfaces  are  placed  over  one  another  with  a  consid- 
erable "gap"  or  space  between.  The  body  is  placed  be- 
tween the  wings  and  the  tail  surfaces  and  chassis  remain 
the  same  as  in  the  monoplane.  This  is  known  as  a  "Trac- 
tor" biplane  since  the  propeller  is  in  front  and  pulls  the 
machine  along  while  Fig.  6  shows  a  "Pusher"  type  biplane 
in  which  the  propeller  is  mounted  behind  the  wings  and 
therefore  pushes  the  machine. 


Fig.  4-A.    Dcperdussin  Monoplane  with  Monocoque  Body.    Gordon-Bennett  Racer 

Biplanes.  Besides  the  advantages  of  size,  the  l)iplane 
has  a  number  of  other  good  features.  The  deep  spacing 
of  the  upper  and  lower  surfaces  permits  of  a  powerful 
and  light  system  of  trussing  being  placed  in  the  gap,  and 
therefore  the  biplane  can  be  made  stronger  (weight  for 
weight)  than  the  monoplane  in  which  no  such  trussing 
can  be  economically  applied.  The  vertical  ''struts"  of  the 
bracing  can  be  clearly  seen  in  the  figure.  The  efficiency 
of  this  interplane  trussing  greatly  increases  the  possible 
size  and  capacity  of  the  aeroplane.  With  monoplanes  it 
is  seldom  possible  to  exceed  a  w^ing  span  of  Z6  feet  with- 
out runing  into  almost  unsurmountable  structural  diffi- 
culties.   The  weight  of  the  large  monoplane  also  increases 


16 


AEROPLANE  PRINCIPLES 


in  leaps  and  bounds  when  this  critical  span  is  once 
exceeded.  To  maintain  an  equal  degree  of  strength  the 
monoplane  requires  very  careful  attention  in  regard  to 
the  design  and  construction,  and  is  correspondingly  more 
expensive  and  difficult  to  build  than  the  biplane,  although 
the  latter  has  by  far  the  greater  number  of  parts.  By 
suitable  arrangements  in  the  location  of  the  biplane  sur- 
faces a  very  fair  degree  of  stability  can  be  obtained,  an 
advantage  which  is  impossible  with  the  monoplane. 


Fig.   5.     S.    P.   A.    D.   Tractor   Biplane   Speed    Scout. 

A  distinct  disadvantage  of  the  two  superposed  sur- 
faces of  the  biplane  is  due  to  the  fact  that  there  is  "inter- 
ference" between  the  upper  and  lower  wings,  and  that  the 
lift  for  equal  areas  is  less  than  in  the  case  of  the  mono- 
plane. With  a  given  form  of  wing,  100  square  feet  of 
monoplane  surface  will  lift  considerably  more  than  the 
same  area  applied  in  biplane  form.  The  amount  of  the 
"drag"  for  the  support  of  a  given  load  is  increased,  and 
with  it  the  amount  of  power  required.  The  greater 
the  separation  or  "gap"  between  the  wings,  the  greater 
will  be  the  lift,  but  when  the  gap  is  unduly  increased  to 
obtain  a  great  lift  the  length  of  the  interplane  bracing  is 
increased  to  such  an  extent  that  the  resistance  of  the 
bracing  will  more  than  overcome  the  advantages  due  to 


AEROPLANE  PRINCIPLES 


17 


the  large  gap.  There  is  a  fixed  limit  to  the  gap  beyond 
which  it  is  not  practical  to  go.  The  bracing  has  a  very 
material  effect  on  the  air  resistance,  no  matter  how  small 
the  gap. 

Triplanes.  Of  late  the  triplane  has  been  rapidly 
increasing  in  use,  and  in  certain  respects  has  many  advan- 
tages over  either  the  monoplane  or  biplane.  This  type 
has  three  superposed  surfaces  which  still  further  dimin- 
ishes the  size  for  a  given  area.    The  interference  between 


Fig.  6.      Pusher  Type  Biplane  in  Which  the  Propeller  Is  Placed  Behind  the  Wings. 

the  surfaces  is  even  greater  than  with  the  biplane,  and 
hence  the  lift  is  less  for  a  given  area  and  aspect  ratio. 
This  latter  defect  is  partly,  or  wholly  overcome  by  the 
possibility  of  using  long  narrow  wings,  and  because  of 
the  reduced  span  there  is  a  corresponding  reduction  in 
the  bracing  resistance.  It  should  be  noted  at  this  point 
that  the  efficiency  of  a  lifting  surface  is  greatly  increased 
when  the  ratio  of  the  length  to  the  width  is  increased, 
that  is,  a  long,  narrow  wing  will  be  more  efficient  than  a 
short,  wide  shape.  The  relation  of  the  length  to  the  width 
is  called  "aspect  ratio,"  and  will  be  described  in  more 
detail  in  a  following  chapter. 

Fig.  7  is  a  sketch  of  a  Sopwith  Triplane  Scout  and 
shows  clearly  the  three  superposed  wings.  The  small 
amount  of  interplane  bracing,  and  the  great  aspect  ratio, 


18 


AEROPLANE  PRINCIPLES 


makes  this  type  very  suitable  for  high  speed.  The  body, 
tail  and  chassis  arrangements  are  practically  the  same  as 
those  of  a  biplane.  The  Curtiss  Triplane  Scout  is  the 
pioneer  of  this  type  of  machine,  although  experimental 


Fig.   6-A.      Farman    Type   Pusher   Biplane.      Note   the    Propeller   At   the   Rear   of 
Body,  and  the  Position  of  the  Pilot  and  Passenger. 


"*■ 

■ 

\ 

« 

-"^w. 

r^j^^^^^WBB(B|^^^ 

^ 

y% 

^?W'  '  ^  f.>V-y '^•^.'/4'     rr^'''  ".    'i 

'^       :i 

ml  l^^^^^^Kt 

Fig.  6-B.  The  Mann  Two-Propeller  Pusher  Biplane.  The  Propellers  Are 
Mountetd  on  Either  Side  of  the  Body,  and  Are  Driven  by  a  Single 
Motor  Through  a  Chain  Transmission.  This  Drive  Is  Similar  to 
the  Early  Wright  Machines. 


AEROPLANE  PRIX'CIPLES 


19 


work  on  the  triplane  had  been  performed  in  England  by 
A.  V.  Roe  many  years  ago.  The  Roe  triplane  was  lightly 
powered  and  for  its  time  was  successful  in  a  way,  but  the 
Curtiss  is  the  first  to  enter  into  active  competition  against 
the  biplane  scout. 

Owing  to  the  small  span  required  for  a  given  area,  and 
the  possibilities  of  very  light  and  simple  bracing,  the 
triplane  is  an  ideal  type  for  heavy  duty  machines  of  the 
"bombing"  species.  Enormous  triplanes  have  been  made 
that  are  capable  of  a  useful  load  running  up  into  the  tons, 


Fig.   7.     Sopwilh    Triplane    Speed    Scout. 


the  large  Curtiss  and  Caproni's  being  notable  examples. 
As  the  triplane  is  much  higher  than  the  biplane  of  equal 
area,  the  interplane  bracing  is  deeper  and  more  effective 
without  causing  proportionately  higher  resistance. 

Quadraplane.  The  use  of  four  superposed  surfaces  has 
not  l)een  extended,  there  probably  being  only  one  or  two 
of  these  machines  that  can  be  said  to  be  successful.  The 
small  "quad"  built  by  Mathew  B.  Sellers  is  probably  the 
best  known.  The  power  required  to  maintain  this  ma- 
chine in  flight  was  surprisingly  small,  the  machine  getting 
off  the  ground  with  a  4  horsepow^er  motor,  although  an 
8  horsepower  was  afterwards  installed  to  maintain  con- 
tinuous flight.  The  empty  weight  was  110  pounds  with 
the  8  horsepower  motor.     The  span  of  the   wings  was 


20  AEROPLANE  PRINCIPLES 

18'— 0"  and  the  width  or  "chord"  3'— 0",  giving  a  total 
area  of  about  200  square  feet. 

Tandem  Aeroplanes.  A  tandem  aeroplane  may  be 
described  as  being  one  in  which  the  surfaces  are  arranged 
*'fore  and  aft."  The  Langley  "Aerodrome"  was  of  the 
tandem  monoplane  type  and  consisted  of  two  sets  of 
monoplane  wings  arranged  in  tandem.  This  pioneer 
machine  is  shown  in  Fig.  8,  and  is  the  first  power  driven 
model  to  achieve  a  continuous  flight  of  any  length. 
Instead  of  two  monoplane  surfaces,  two  biplane  units  or 


Fig.   7-A.     Curtiss   Triplane   Speed    Scout.      Courtesy   "Aerial  Age." 

triplane  units  can  be  arranged  fore  and  aft  in  the  same 
manner. 

While  there  have  been  a  number  of  tandem  machines 
built,  they  have  not  come  into  extensive  use.  Successful 
flight  was  obtained  with  a  full  size  Langley  Aerodrome, 
and  this  machine  flew  with  a  fair  degree  of  stability.  The 
failure  of  other  tandem  machines  to  make  good  was  due, 
in  the  writer's  opinion,  to  poor  construction  and  design 
rather  than  to  a  failure  of  the  tandem  principle.  The 
Montgomery  glider,  famed  for  its  stability,  was  a  tandem 
type  but  the  machine  was  never  successfully  built  as  a 
powered  machine. 

The  wings  must  be  separated  by  a  sufficient  distance  so 


AEROPLANE  PRINXIPLES  21 

that  the  rear  set  will  not  be  greatly  influenced  by  the 
downward  trend  of  the  air  caused  by  the  leading  wings. 
As  the  rear  surfaces  always  work  on  disturbed  air  they 
should  be  changed  in  angle,  increased  in  area,  or  be 
equipped  with  a  different  wing  curvature  if  they  are  to 
carry  an  equal  proportion  of  the  load.  Usually,  however^ 
the  areas  of  the  front  and  rear  wings  are  equal,  and  the 
difference  in  lift  is  made  by  changes  in  the  wing  form  or 
angle  at  which  they  are  set.  In  some  cases  the  wings  are 
approximately  the  same,  the  difference  in  lift  being  com- 
pensated for  by  moving  the  load  further  forward,  thus, 
throwing  more  of  the  weight  on  the  front  wings. 

The  Aeroplane  in  Flight.  Up  to  the  present  we  have 
only  considered  horizontal  flight  at  a  continuous  speed. 
In  actual  flight  the  altitude  is  frequently  varied  and  the 
speed  is  changed  to  meet  different  conditions.  Again,  the 
load  is  not  an  absolutely  constant  quantity  owing  to 
variations  in  the  weight  of  passengers,  and  variations  in 
the  weight  of  fuel,  the  weight  of  the  latter  diminishing 
directly  with  the  length  of  the  time  of  flight.  To  meet 
these  variations,  the  lift  of  the  wings  must  be  altered  to 
suit  the  loading  and  speed — generally  by  altering  the 
angle  of  the  wings  made  with  the  line  of  flight. 

Fig,  9  shows  an  aeroplane  in  horizontal  flight  and  lightly^ 
loaded,  the  machine  traveling  along  the  horizontal  flight 
path  F-F.  With  the  light  load,  the  angle  made  by  the 
wings  with  the  flight  path  is  shown  by(i),  the  tail  and  body 
remaining  horizontal,  or  parallel  to  the  flight  path.  With 
an  increased  load  it  is  necessary  to  increase  the  angle  of 
the  wings  with  the  flight  line,  since  within  certain  limits 
the  lift  increases  with  an  increase  in  the  angle  of  incidence 
(i).  Fig.  10  shows  the  adjustment  for  a  heavier  load  (Wo)^ 
the  angle  of  incidence  being  increased  to  (i'),  and  the 
body  is  turned  down  through  a  corresponding  angle.  The 
increased  angle  is  obtained  by  turning  the  elevator  flaps 
(T)  up,  thus  causing  a  downward  force  (t)  on  the  tail. 


22 


AEROPLANE  PRINCIPLES 


The  force  (t)  acts  through  the  body  as  a  lever  arm,  and 
turns  the  machine  into  its  new  position.  It  will  be  noted 
that  when  the  angle  of  incidence  is  great  that  the  rear  of 
the  body  drags  down  and  causes  a  heavy  resistance.  This 
position  of  a  dragging  tail  is  known  to  the  French  as 
flying  *'Cabre."  With  high  angles  cabre  -flight  is  dan- 
gerous, for  should  the  propeller  thrust  cease  for  an 
instant  the  machine  would  be  likely  to  "tail  dive"  before 


T"ig.   8.      Langley's  "Aerodrome,"  An  Early  Type  of  Tandem   Monoplane. 


the  pilot  could  regain  control.  This  sort  of  flight  is  also 
wasteful  of  power.  Cabre  flight  is  unnecessary  in  a 
variable  incidence  machine,  the  wing  being  turned  to  the 
required  angle  independently  of  the  body,  so  that  the 
body  follows  the  flight  line  in  a  horizontal  position,  no 
njatter  what  the  angle  of  incidence  may  be.  In  this  type 
of  machine  the  wings  are  pivoted  to  the  body,  and  are 
operated  by  some  form  of  manual  control. 


AEROPLANE  PRINCIPLES 


23 


In  Fig.  11,  the  large  angle  (i')  is  still  maintained,  but 
the  load  is  reduced  to  the  value  given  in  Fig.  9.  With  an 
equal  load,  an  increased  angle  of  incidence  causes  the 
machine  to  climb,  as  along  the  new  flight  line  f-f.  With 
the  load  ( W)  equal  to  that  in  Fig.  9,  the  angle  of  incidence 
Avill  still  be  (i)  but  this  will  be  along  a  new  flight  line  if 
the  large  angle  (i')  is  maintained  with  the  horizontal  as 


tICAvy  LO^D-SLjaW  SfEED 


Flight  At  t:  rimEa 
Que  re  CHAf^ees  "^  SP££D.  kve/e^rr^n 

CM  A  VS;?S  IN  /IL  TIT  LUX. .  


Figs.  9-10-11-12.  Showing  the  Use  of  Elevators  in  Changing  Angle  of  Incidence. 
Machine  Shown  in  Four  Principal  Attitudes  of  Flight.  As 
l.^^  J?°^^>'  ^"'^  Wings  Are  in  a  Single  Unit,  the  Body  Must 
Be  Turned  for  Each  Different  Wing  Angle. 


shown  by  Fig.  11.  With  the  wings  making  an  angle  of 
(i')  with  the  horizontal,  and  angle  of  incidence  (i)  with 
the  flight  line,  it  is  evident  from  Fig.  11  that  the  new 
flight  line  f-f  must  make  an  angle  (c)  with  the  original 
horizontal  flight  line  F-F.  This  shows  how  an  increased 
angle  with  a  constant  load  causes  climbing,  providing,  of 
course,  that  the  speed  and  power  are  maintained.  With 
a  given  wing  and  load  there  is  a  definite  angle  of  incidence 
if  the  speed  is  kept  constant.  Should  a  load  be  dropped, 
such  as  a  bomb,  with  the  wing  angle  kept  constant,  the 
new  path  of  travel  will  be  changed  from  F-F  to  f-f. 

Fig.  12  shows  the  condition  when  the  rear  end  of  the 
body  is  elevated  by  depressing  the  elevator  flap  T.    This 


24  AEROPLANE  PRINCIPLES 

occasions  an  upward  tail  force  that  turns  the  wings  down 
through  the  total  angle  (i').  With  the  former  loading 
and  speed,  the  angle  of  incidence  is  still  (i)  degrees  with 
the  new  flight  path  f-f,  the  new  flight  path  being  at  an 
angle  (c)  with  the  horizontal  F-F.  The  body  is  turned 
through  angle  (i'),  but  the  angle  (i)  with  the  flight  path 
f-f  is  still  constant  with  equal  loads  and  speeds. 

To  cause  an  aeroplane  to  climb,  or  to  carry  a  heavier 
load,  the  elevator  *'flap"  is  pointed  up.  To  descend,  or 
care  for  a  lighter  load,  the  elevator  is  turned  dow^n.  In 
normal  horizontal  flight  the  machine  should  be  balanced 
so  that  the  tail  is  horizontal  and  thus  creates  no  drag. 
When  the  elevator  must  be  used  to  keep  the  tail  up  in 
horizontal  flight,  the  machine  is  said  to  be  "tail  heavy." 

Longitudinal  Stability.  In  Figs.  9-10-11-12  the  machine 
was  assumed  to  be  flying  in  still  air,  the  attitudes  of  the 
machine  being  simply  due  to  changes  in  the  loading  or  to 
a  change  in  altitude.  The  actual  case  is  more  complicated 
than  this,  for  the  reason  that  the  machine  is  never  operat- 
ing in  still  air  but  encounters  sudden  gusts,  whorls,  and 
other  erratic  variations  in  the  density  and  velocity  of  the 
air.  Each  variation  in  the  surrounding  air  causes  a  change 
in  the  lift  of  the  wings,  or  in  the  effect  of  the  tail  surfaces, 
and  hence  tends  to  upset  the  machine.  If  such  wind  gusts 
would  always  strike  the  w^ngs,  body,  and  tail  simul- 
taneously, there  would  be  no  trouble,  but,  unfortunately, 
the  air  gust  strikes  one  portion  of  the  machine  and  an 
appreciable  length  of  time  elapses  before  it  travels  far 
enough  to  strike  another.  Though  this  may  seem  to  be 
a  small  fraction  of  time,  it  is  in  reality  of  sufficient  dura- 
tion to  have  a  material  effect  on  the  poise  of  the  aero- 
plane. Vertical  gusts  due  to  the  wind  passing  over  build- 
ings, hills,  cliffs,  etc.,  not  only  tend  to  upset  the  machine, 
but  also  tend  to  change  the  altitude  since  the  machine 
rises  with  an  up  gust  and  sinks  with  a  down  trend  in  the 
stream. 


AEROPLANE  PRINXIPLES  25 

Assume  a  machine  as  in  Fig.  9  to  be  traveling  steadily 
along  a  horizontal  path  in  still  air.  A  sudden  horizontal 
gust  now  strikes  the  machine  from  the  front,  thus  causing 
a  sudden  lift  in  the  main  wings.  As  this  gust  strikes  the 
wings  before  the  tail,  the  tail  will  stand  at  the  old  altitude 
while  the  wings  are  lifted,  thus  giving  the  position  shown 
by  Fig.  10.  After  passing  over  the  wings  it  lifts  the  tail, 
this  effect  probably  not  being  sufficient  to  restore  the 
wing  and  the  tail  to  their  old  relative  attitude  since  the 
gust  generally  loses  velocity  after  passing  the  wings.  A 
head  gust  of  this  type  often  strikes  the  front  wings  diag- 
onally so  that  it  never  reaches  the  tail  at  all.  To  remedy 
this  upsetting  action  of  the  gust,  the  pilot  must  move  his 
rear  elevator  so  that  the  elevator  is  in  the  position  shown 
by  Fig.  12,  that  is,  the  flap  must  be  turned  down  so  as  to 
raise  the  tail. 

A  gust  striking  from  behind  may,  or  may  not  afifect  the 
elevator  flaps,  this  depending  on  their  position  at  the  time 
that  the  gust  strikes.  If  the  flaps  are  turned  up,  the  rear 
end  will  be  raised  by  the  gust  and  the  machine  will  head 
dive :  if  turned  down,  the  gust  will  depress  the  tail,  raise 
the  head  and  tend  to  ''stall"  the  machine.  If  the  tail  is  of 
the  lifting  type,  the  rear  entering  gust  will  reduce  the 
relative  velocity,  and  the  lift,  and  cause  the  tail  to  drop.  On 
passing  over  the  tail  and  striking  the  wings,  the  rear 
gust  will  reduce  the  velocity  and  cause  a  loss  in  lift. 
This  will  either  cause  the  machin.e  to  head  dive  or  drop 
vertically  through  a  certain  distance  until  it  again  assumes 
its  normal  velocity. 

All  of  these  variations  cause  a  continually  fore  and  aft 
upsetting  movement  that  must  be  continually  corrected 
by  raising  and  lowering  the  elevator  flaps,  and  in  very 
gusty  weather  this  is  a  very  tedious  and  wearing  job.  It 
requires  the  continual  attention  of  the  pilot  unless  the 
action  is  perforrned  automatically  by  some  mechanical 
device,  such  as  the  Sperry  Gyroscopic,  or  else  by  some 


26  AEROPLANE  PRINCIPLES 

arrangement  of  the  surfaces  that  give  ''inherent"  stabil- 
ity. Control  by  means  of  the  elevator  flaps  (v^^hich  raise 
and  lower  the  body  in  a  fore  and  aft  direction,  as  shown) 
is  known  as  "longitudinal  control,"  and  when  the  machine 
is  so  built  that  correction  for  the  longitudinal  attitude  is 
obtained  ''inherently,"  the  machine  is  said  to  be  "longi- 
tudinally stable."  Modern  machines  can  be  made  very 
nearly  longitudinally  stable,  and  are  comparatively  un- 
affected by  any  than  the  heaviest  gusts. 

Lateral  Stability.  The  gusts  also  affect  the  side  to  side, 
or  "lateral"  balance  by  causing  a  dift'erence  in  lift  on  either 
end  of  the  wings.  Should  the  gust  strike  one  tip  before 
the  other,  or  should  it  strike  one  tip  harder  than  the 
other,  the  tendency  will  be  to  turn  the  machine  over  side- 
wise.  This  is  a  more  difficult  problem  to  solve  than  the 
longitudinal  moment,  although  perfect  inherent  stability 
has  been  attained  in  one  or  two  machines  without  the 
use  of  additional  automatic  control  mechanism.  Inherent 
lateral  stability  has  always  been  attended  by  a  considerable 
loss  in  the  efficiency  of  the  aeroplane  and  speed  due  to  the 
peculiar  arrangements  in  the  main  lifting  surfaces.  At 
present  we  must  make  a  decision  between  efficiency  and 
stability,  for  one  feature  must  be  attained  at  a  sacrifice 
in  the  other.  Contrary  to  the  general  opinion,  perfect 
stability  is  not  desirable,  for  almost  invariably  it  affects 
the  control  of  a  machine  and  makes  it  difficult  to 
maneuver.  Should  the  stability  appliances  be  arranged 
so  that  they  can  be  cut  out  of  action  at  Avill,  as  in  the  case 
of  the  Sperry  Gyroscopic  Stabilizer,  they  will  fulfill  the 
needs  of  the  aviator  much  more  fully  than  those  of  the 
fixed  inherent  type.  The  first  thoroughly  stable  machine, 
both  longitudinally  and  laterally,  was  that  designed  by 
Lieutenant  Dunne,  and  this  obtained  its  distinctive  fea- 
ture by  a  very  peculiar  arrangement  of  the  wing  surfaces. 
It  was  excessively  stable,  and  as  with  all  very  stable 
machines,   was   difficult   to   steer   in   a   straight   line   in 


AEROPLANE  PRIXXIPLES 


28  AEROPLANE  PRINCIPLES 

windy  weather,  and  was  correspondingly  difficult  to 
land. 

The  first  machine  of  the  ordinary  biplane  type  that 
proved  inherently  stable  was  the  R.  E.-l  designed  in 
England  by  Edward  Busk.  This  machine  was  flown  from 
Farnborough  to  Salisbury  Plain,  and  during  this  flight  the 
only  control  touched  was  the  vertical  rudder  used  in  steer- 
ing. Since  then,  all  English  machines  have  been  made  at 
least  partially  stable,  the  degree  depending  upon  the  serv- 
ice for  which  it  was  intended.  It  has  been  found  that  in 
lighting,  a  very  controllable  machine  is  necessary,  hence 
stability  must  be  sacrificed,  or  the  control  surfaces  must 
be  made  sufficiently  powerful  to  overcome  the  stable  tend- 
ency of  the  machine.  War  machines  are  made  to  be  just 
comfortably  stable  over  the  range  of  ordinary  flight 
speeds,  and  with  controls  powerful  enough  so  that  the 
inherent  stability  can  be  overcome  when  maneuvering 
in  battle.  The  present  war  machine  always  contains  an 
element  of  danger  for  the  unskilled  pilot. 

Dihedral  Angle.  This  was  the  first  lateral  stability 
arrangement  to  be  applied  to  an  aeroplane,  but  is  only 
effective  in  still  air.  In  rough  weather  its  general  tend- 
ency is  to  destroy  stability  by  allowing  dangerous  oscilla- 
tions to  take  place.  Fig.  13  is  a  front  view  of  a  monoplane 
in  which  the  wings  (w)  and  (w')  are  set  at  an  angle  (d), 
this  angle  being  known  as  the  "dihedral  angle."  The 
dotted  line  (m-m)  shows  the  line  of  a  pair  of  perfectly 
horizontal  wings  and  aids  in  illustrating  the  dihedral. 
Assuming  the  center  of  lift  at  CL  on  the  wings,  it  will 
be  seen  that  an  increase  in  the  dihedral  raises  the  center 
•of  lift  above  the  center  of  gravity  line  C.  G.  by  the  amount 
(h).  With  the  center  of  gravity  below  the  center  of  lift 
it  is  evident  that  the  weight  of  the  machine  would  tend 
to  keep  it  on  a  level  keel,  although  the  same  effect  could, 
of  course,  be  attained  in  another  way.  The  principal 
righting  effect  of  the  dihedral  is  shown  by  Fig.   14  in 


AEROPLANE  PRINXIPLES 


29 


which  the  wings  (w)  and  (w')  are  set  as  before.  The 
machine  is  tipped  or  "Hsted"  toward  the  left  (seen  from 
aviator's  seat)  so  that  wing  (w')  is  down.  By  drawing 
vertical  lines  down  until  they  intersect  the  horizontal  line 
X-X  (the  line  of  equilibrium),  it  will  be  seen  that  wing 
(w')  presents  more  horizontal  lift  surface  than  (w)  since 
the  projected  or  effective  wing  length  (C)  is  greater  than 
(b).  Since  (w')  presents  the  greater  surface,  the  lift  (L) 
tends  to  restore  the  machine  to  its  original  level  position. 
If  the  wings  were  both  set  on  the  same  straight  line,  the 


Fig.  1 5 


naie 


projected  lengths   (b)   and   (c)  would  be  the  same  and 
there  would  be  no  restoring  effect. 

The  dihedral  would  be  very  effective  in  still  air,  but  in 
turbulent  air,  and  with  the  body  swinging  back  and  forth, 
the  dihedral  would  act  in  the  nature  of  a  circular  guiding 
path,  and  thus  tend  to  allow  the  swinging  to  persist  or 
increase  rather  than  to  damp  it  down,  as  would  be  the 
case  with  level  straight  wings.  Again,  with  the  wing  bent 
up  at  a  considerable  angle,  a  side  gust  as  at  (S)  would 
tend  to  throw  the  machine  still  further  over,  and  thus 


30  AEROPLANE  PRINCIPLES 

increase  rather  than  diminish  the  difficulty.  In  prac- 
tical machines,  the  dihedral  is  usually  made  very  small 
(d=  176  degrees),  the  angle  of  each  wing  with  the  hori- 
zontal being  about  2  degrees  or  even  less.  I  think  the 
advantage  of  such  a  small  angle  is  rather  more  imaginary 
than  actual,  and  at  present  the  greater  number  of  war 
machines  have  no  dihedral  at  all.  In  the  older  monoplanes 
the  angle  was  very  pronounced. 

Fig.  15  shows  the  dihedral  applied  both  to  the  upper 
wing  (U)  and  the  lower  wing  (L),  the  usual  method  of 
applying  dihedral  to  large  biplanes.  Fig.  16  shows  the 
method  of  applying  the  dihedral  to  small,  fast  machines, 
such  as  speed  scouts,  the  dihedral  in  this  case  being  used 
only  on  the  lower  wing.  The  dihedral  on  the  bottom  wing 
is  usually  for  the  purpose  of  clearing  the  wing  tips  when 
turning  on  the  ground  rather  than  for  stability.  A  loAver 
wing  with  a  dihedral  is  less  likely  to  strike  the  ground  or 
to  become  fouled  when  it  encounters  a  side  gust  in  land- 
ing or  ''getting  off.''  The  use  of  straight  upper  wings 
makes  the  construction  much  simpler,  especially  on  the 
small  machines  where  it  is  possible  to  make  the  wing 
in  one  continuous  length. 

Ailerons  and  Wing  Warping.  Since  the  dihedral  is  not 
effective  in  producing  lateral  stability,  some  other  method 
must  be  used  that  is  powerful  enough  to  overcome  both 
the  upsetting  movements  and  the  lateral  oscillations 
caused  by  the  pendulum  effect  of  a  low  center  of  gravity. 
In  the  ordinary  type  of  aeroplane  this  righting  effect  is  per- 
formed by  movable  surfaces  that  increase  the  lift  on  the 
lower  Aving  tip,  and  decrease  the  lift  on  the  high  side.  In 
some  cases  the  lateral  control  surfaces  are  separate  from 
the  wing  proper  (Ailerons),  and  in  some  the  tip  of  the 
wing  is  twisted  or  "warped"  so  as  to  produce  the  same 
effect.  These  control  surfaces  may  be  operated  manually 
by  the  pilot  or  by  some  type  of  mechanism,  such  as  the 
gyroscope,  although  the  former  is  the  method  most  used. 


AEROPLANE  PRINCIPLES  31 

The  lateral  control,  or  side  to  side  balancing  of  an  aero- 
plane, can  be  compared  to  the  side  to  side  balancing  of 
a  bicycle  in  which  the  unbalance  is  continually  being 
corrected  by  the  movement  of  the  handle  bars. 

Fig.  17  shows  the  control  surfaces  or  "Ailerons" 
(A-A'),  mounted  near  the  tips,  and  at  the  rear  edge  of  the 
wing  W.  As  shown,  they  are  cut  into  and  hinged  to  the 
main  wings  so  that  they  are  free  to  move  up  and  down 
through  a  total  angle  of  about  60  degrees.  In  a  biplane  they 
may  be  fitted  to  the  upper  wing  alone  or  to  both  top  and 
bottom  wings,  according  to  conditions.  For  simplicity  we 
will  consider  only  the  monoplane  in  the  present  instance. 

In  Fig.  18,  a  front  view  of  the  monoplane,  the  machine 
is  shown  "heeled  over"  so  that  the  wing  tip  (w)  is  low. 
To  correct  this  displacement,  the  aileron  (A)  on  the  low 
side,  is  pulled  down  and  the  aileron  (A')  on  the  opposite 
end  is  pulled  up.  This,  of  course,  increases  the  lift  on 
the  low  end  (w)  and  decreases  the  lift  on  the  high  side 
(w').  The  righting  forces  exerted  are  shown  by  L-L'. 
The  increased  angles  made  by  A-A'  with  the  wind  stream 
affects  the  forces  acting  on  the  wings,  although  in  oppo- 
site directions,  causing  a  left  hand  rotation  of  the  whole 
machine.  In  Fig.  19,  conditions  are  normal  with  the  ma- 
chine on  an  even  keel  and  with  both  ailerons  brought  back 
to  a  point  where  they  are  level  with  the  surface  of  the 
wing,  or  in  "neutral."  Fig.  20  shows  the  machine  canted 
in  the  opposite  direction  with  (w')  low  and  (w)  high. 
This  is  corrected  by  bringing  down  aileron  (A')  and  rais- 
ing (A),  the  forces  L-L'  showing  the  rotation  direction. 
By  alternately  raising  and  lowering  the  ailerons  we  can 
correspondingly  raise  or  lower  the  wing  tips.  It  should 
be  noted  here  that  in  some  machines  the  ailerons  are  only 
ingle  acting,  that  is,  the  aileron  on  the  low  side  can  be 
pulled  down  to  increase  the  lift,  but  the  opposite  aileron 
remains  in  the  plane  of  the  wings,  and  does  not  tend  to 
"push  down"  the  high  side.    Since  all  of  the  aileron  resist- 


32  AEROPLANE  PRINCIPLES 

ance  in  a  horizontal  direction  is  now  confined  to  the  low 
side,  it  turns  the  machine  from  its  path,  the  high  wing 
swinging  about  the  lower  tip  with  the  latter  as  a  pivot. 
In  the  double  acting  control  as  shown  in  Figs.  17  to  23, 
the  resistance  is  nearly  equal  at  both  tips  and  hence  there 
is  no  tendency  to  disturb  the  flight  direction.  With  single 
acting  ailerons,  the  directional  disturbance  is  corrected 
with  the  rudder  so  that  when  the  aileron  is  pulled  down 
it  is  necessary  to  set  the  rudder  to  oppose  the  turn.  On 
early  Wright  machines  the  rudder  and  lateral  controls 
were  interconnected  so  that  the  rudder  automatically 
responded  to  the  action  of  the  ailerons. 

Fig.  21  is  a  detailed  front  elevation  of  the  machine  and 
shows  the  control  wheel  (C)  and  cable  connections 
between  the  wheel  and  the  ailerons  A-A'.  When  the 
wheel  C  is  turned  in  the  direction  of  the  arrow  K,  the 
aileron  A'  is  pulled  down  by  the  flexible  cable  (i),  and  a 
corresponding  amount  of  cable  (h)  is  paid  off  the  wheel  to 
the  rising  aileron  A.  Aileron  A  is  pulled  up  by  the  con- 
necting cable  (e)  which  is  attached  to  A'  at  one  end  and 
to  A  at  the  other.  Pulleys  (f)  and  (g)  guide  the  intercon- 
nection cable.  On  turning  the  wheel  in  the  opposite  direc- 
tion, aileron  A  is  pulled  down  and  A'  is  elevated.  In 
flight,  especially  in  rough  weather,  there  is  almost  con- 
tinuous operation  of  the  control  wheel.  Figs.  22  and  23 
are  sections  taken  through  the  wing  W  and  the  ailerons, 
showing  the  method  of  hinging  and  travel.  Fig.  22  shows 
the  aileron  depressed  for  raising  the  wing  in  the  direction 
of  L,  while  Fig.  23  shows  the  aileron  lifted  to  lower  the 
wing.  In  normal  flight,  with  the  machine  level,  the  aileron 
forms  a  part  of  the  wing  outline  (in  neutral  position). 

In  the  original  Wright  aeroplane,  and  in  the  majority  of 
monoplanes,  no  ailerons  are  used,  the  rear  of  main  wing 
tip  being  bent  down  bodily  to  increase  the  lift.  This  is 
known  as  "wing  warping,"  and  is  practically  a  single 
acting  process  since  the  depressing  force  on  the  high  tip 


AEROPLANE  PRINXIPLES 


33 


is  seldom  as  effective  as  the  lift  on  the  low.  Warping  is 
not  generally  used  on  modern  biplanes  since  it  is  impos- 
sible to  maintain  a  strong  rigid  structure  with  flexible 
tips.  The  control  warping  and  twisting  of  such  wings 
soon  loosens  them  up  and  destroys  what  remaining 
strength  they  may  have  had. 


FlG-gg 


Fis^ga 


Figs.    17-23.     Showing  Use   of  Ailerons  in  Maintaining  Lateral   Balance. 


Banking  and  Turning.  In  making  a  sharp  turn  the 
outer  wing  tip  must  be  elevated  to  prevent  slipping  side- 
wise  through  the  effects  of  the  centrifugal  force  (side 
slip).  This  is  known  as  "banking."  The  faster  and 
sharper  the  turn,  the  steeper  must  be  the  "bank,"  or  the 
angle  of  the  wings,  until  in  some  cases  of  "stunt"  flying 
the  wings  stand  almost  straight  up  and  down.    Should  the 


34  AEROPLANE  PRINCIPLES 

bank  be  too  steep  there  will  be  an  equal  tendency  to  slip 
down,  and  inwardly,  since*  the  end  resistance  against  side 
slip  is  very  slight.  Some  aeroplanes  assume  the  correct 
angle  of  bank  automatically  without  attention  from  the 
pilot  since  the  extra  lift  due  to  the  rapid  motion  of  the 
outer  tip  causes  it  to  rise.  On  other  machines  the  natural 
banking  effort  of  the  machine  is  not  sufficient,  and  this 
must  be  increased  by  pulling  down  the  aileron  on  the  outer 
wing  tip.  Machines  that  have  a  tendency  to  "over-bank" 
must  have  the  ailerons  applied  in  the  reverse  direction  so 
as  to  depress  the  outer  tip.  In  cases  of  under,  or  over- 
banking  machines  it  formerly  required  experience  and 
judgment  on  the  part  of  the  pilot  to  obtain  the  correct 
banking  angle.  There  are  now  ''banking  indicators"  on 
the  market  that  show  whether  the  machine  is  correctly 
banked  or  is  side-slipping. 


AEROPLANE  PRINXIPLES 


35 


:Mhs 


Fig.  24.  A  Deperdussin  Monoplane  Banking  Around  a  Sharp  Turn  at  High 
Speed.  Note  the  Elevation  of  the  Outer  Wing  Tip  and  the  Angle 
Made  with  the  Horizontal  by  the  Wings.     Speed,  105  Miles  Per  Hour. 


CHAPTER  II. 
TYPES  OF  MILITARY  AEROPLANES. 

Divisions  of  Service.  In  the  army  and  navy,  aeroplanes 
are  used  both  for  offensive  and  defensive  operations. 
They  must  carry  out  their  own  work  and  intentions  and 
prevent  hostile  craft  from  carrying  out  theirs.  In  of- 
fensive operations  the  machines  fly  continuously*  over 
the  enemy's  country  and  attack  every  hostile  craft  sighted, 
thus  creating  a  danger  zone  within  the  enemy's  lines 
where  no  opposing  machine  can  work  without  being 
threatened  with  an  overwhelming  attack.  The  oft'ensive 
also  includes  bombing  operations  and  the  destruction 
of  supply  depots  and  transportation  centers.  Defensive 
aerial  operation  consists  in  driving  out  the  enemy  craft 
from  our  own  lines,  and  in  protecting  working  machines 
when  on  photographing  or  observation  trips.  With  a 
powerful  offensive  there  is  of  course  little  need  for  de- 
fense. The  former  method  is  a  costly  one,  and  is  pro- 
ductive of  severe  material  and  personal  losses. 

At  the  present  time  there  are  eight  principle  functions 
performed  by  military  aeroplanes : 

1.  Offensive  operations  against  enemy  machines. 

2.  Reconnaissance,  observation,  special  missions. 

3.  Bombing  supply  centers,  railways,  etc. 

4.  Photography. 

5.  Spotting  for  the  artillery. 

6.  Signalling  for  infantry  operations. 

7.  Submarine  hunting. 

8.  Patrol  and  barrage. 

37 


38  MILITARY  AEROPLANES 

Probably  the  most  important  service  of  all  is  per- 
formed by  machines  mider  heading  (1).  If  a  successful 
offensive  can  be  maintained  over  the  enemy's  lines  he 
is  unable  to  intelligently  direct  his  artillery  fire,  and 
can  obtain  no  information  regarding  reinforcements,  or 
troop  concentrations  for  an  impending  attack.  With 
fighting  aeroplanes  clearing  the  way  for  our  own  observa- 
tion machines  and  artillery  spotters,  the  enemy  is  not 
only  blinded,  but  is  blocked  in  any  attempt  to  attack  or 
concentrate  his  forces.    The  fact  that  the  French  aerial 


Fig.    1.     Curtios    "Baby"    Biplane    Speed    Scout.      Equipped    with     100    Horse- 
Power  Water  Cooled  Motor, 


offensive  at  Verdun  was  so  efficiently  and  well  maintained 
accounts  for  the  failure  of  the  heavy  German  artillery. 
Driven  far  back  over  their  own  lines,  the  German  aviators 
were  seldom  able  to  observe  the  placing  of  the  shells, 
and  as  a  result  their  gunners  were  practically  trusting 
to  luck  in  reaching  their  target.  An  immediate  and  accu- 
rate bombardment  always  followed  one  of  the  very  in- 
frequent German  air  raids  over  the  Fren'ch  lines.  When- 
ever the  French,  partially  abandoned  their  aerial  offen- 
sive in  favor  of  a  defensive  campaign,  they  soon  lost 
their  mastery  of  the  air.  As  long  as  enemy  machines  can 


MILITARY  AEROPLANES  39 

be  kept  back  of  their  own  lines,  new  trench  systems  can 
be  constructed,  transportation  lines  can  be  extended  and 
ammunition  dumps  arranged,  undertakings  that  would 
be  highly  precarious  with  enemy  observation  machines 
continually  passing  overhead. 

To  maintain  an  effective  oft'ensive  places  a  tremendous 
strain  on  both  the  men  and  the  machines,  for  though 
the  aeroplanes  do  not  penetrate  far  beyond  the  lines  they 
usually  meet  with  superior  numbers,  and  in  addition  are 
continually   in    range    of   the    anti-aircraft   guns.     In   an 


Fig.   1.     Italian  "Pomilio"  Two  Seater   Biplane.      Courtesy  "Flying." 

attack  over  hostile  country  a  slight  mishap  may  cause 
the  loss  of  a  'plane,  for  usually  the  distance  from  its  base 
is  so  great  as  to  prevent  a  gliding  return.  Over  its  own 
lines  an  engine  failure  is  usually  only  a  temporary  incon- 
venience. 

Fighting  aeroplanes,  for  the  oft'ensive,  are  small  high 
powered  machines  generally  of  the  single  seater  type,  and 
are  capable  of  high  horizontal  and  climbing  speeds.  The 
armament  consists  of  a  machine  gun  of  the  Lewis  type, 
and  occasionally  a  few  light  bombs  may  be  included  in  the 
equipment.    As  they  do  not  carry  out  operations  far  to  the 


40 


MILITARY  AEROPLANES 


rear  of  the  enemy's  lines  they  are  provided  with  fuel  for 
only  two  or  three  hours,  and  this  reduced  fuel  load  is 
necessary  for  the  high  speeds  that  must  necessarily  be 
attained.  The  area  is  limited  to  permit  of  quick  maneuvers 
in  attack  and  escape,  and  at  the  same  time  to  reduce  the 
head  resistance  and  weight.  The  horizontal  speed  may 
run  up  to  150  miles  per  hour,  with  a  climbing  velocity  that 
may  exceed  1,000  feet  per  minute.     Such  machines  are 


Fig.  2.  Machine  Gun  Mounting  on  Morane  Monoplane.  Gun  Fires  Directly 
Through  the  Propeller  Disc.  The  Deflecting  Plate  Attached  to  the 
Root  of  the  Propeller  Blade  Protects  the  Propeller  When  in  Line  of 
Fire.  Ammunition  in  This  Gun  Is  Furnished  in  Straight  Strips  or 
"Clips." 


variously  known  as  "Speed  scouts,"  **Chasers"  or  "Pursuit 
type"  (French  "DeChasse").  At  the  beginning  of  the  war 
the  chasers  were  largely  of  the  monoplane  type,  but  at 
present  the  biplane  is  in  almost  exclusive  use. 

The  aeroplane  employed  for  surveys  of  the  enemy  coun- 
try and  battle  front  (2)  are  of  an  entirely  different  type 
and  are  much  larger  and  slower.  These  "Reconnaissance" 
machines  are  generally  of  the  two-seater  type,  the  per- 


MILITARY  AEROPLANES 


41 


sonnel  consisting  of  an  observer  and  the  pilot,  although  in 
some  cases  a  third  man  is  carried  as  an  assistant  to  the 
observer,  or  to  handle  a  machine  gun  against  an  attack. 
Since  their  speed  is  comparatively  low,  they  are  generally 
provided  with  an  escort  of  chasers,  especially  when  em- 
ployed on  distant  missions,  this  escort  repelling  attacks 
while  the  observations  are  being  made. 

For  accurate  observation  and  mapping,  the  speed  of  an 


Fig.  2-a.     Machine   Gun    Mounting   on    S.    P.    A.    D.    Biplane. 
Attached  to   Fuselage  Top  in  Front  of  Pilot. 


Gun   Is    Rigidly 


observation  machine  must  be  necessarily  low,  and  as  they 
are  additionally  burdened  with  a  wireless  set,  an  observer, 
a  large  fuel  reserve,  and  other  impedimenta,  they  have  a 
comparatively  great  area  and  are  therefore  lacking  in  the 
maneuvering  qualities  of  the  chaser.  The  span  will  aver- 
age about  40  feet,  and  the  weight  carried  per  horsepower 
is  greatly  in  excess  of  that  of  the  chaser.  From  a  number 
of  examples,  the  reconnaissance  type  will  average  from 
16  to  18  pounds  per  horsepower,  while  the  loading  of  the 
scout  is  from  8  to  12.  This  means  that  the  former  has  com- 


42  MILITARY  AEROPLANES 

paratively  little  reserve  power  for  rapid  climbing.  The 
present  reconnaissance  type  is  always  armed,  and  must 
not  be  confused  with  the  early  machine  by  that  name, 
which,  in  fact,  was  merely  an  enlarged  training  machine 
and  had  neither  offensive  nor  defensive  powers.  The  ob- 
server acts  as  gunner,  and  is  located  at  a  point  where  he 
has  the  greatest  possible  range  of  vision,  and  where  the 
angle  of  fire  is  as  little  obstructed  as  possible. 

The  radius  of  action,  or  the  distance  traveled  per  tank 
of  fuel,  is  greater  with  the  reconnaissance  than  with  the 
chaser,  present  machines  having  a  capacity  of  from  10  to 
12  hours  on  a  single  filling  at  normal  flight  speed. 

In  bombing  operations  (3),  the  loading  is  very  heavy 
and  consequently  a  ''Bomber"  must  be  a  weight  lifter  to 
the  exclusion  of  all  other  qualities.  Not  only  is  the  bomb 
load  requirement  severe,  but  the  fuel  load  is  also  of  great 
importance,  since  bombing  is  usually  carried  out  at  con- 
siderable distances  from  the  base.  Such  machines  may 
carry  from  three  to  six  men.  All  this  calls  for  a  tremen- 
dous area  and  a  large  power  plant.  The  Handley-Page 
*'Giant,"  and  the  Caproni  Triplane  a^e  examples  of  Allied 
machines  ot  this  type  w^hile  the  German  "Gotha,"  used 
in  the  London  air  raids,  is  an  equi\^alent  enemy  machine. 
As  an  example  of  the  weight  carrying  capacity  of  a 
typical  bomber,  the  Handley-Page  has  carried  a  test  crew 
of  21  men,  or  a  personnel  load  of  3,570  pounds.  The  total 
weight,  fully  loaded,  has  been  given  as  11,500  pounds 
with  a  power  plant  of  540  horsepower.  The  maximum 
speed  is  90  miles  per  hour  with  a  climbing  velocity  of 
about  330  feet  per  minute.  Duration  is  about  5^  hours 
at  normal  speed  and  full  load. 

Bombing  is  of  great  importance,  not  only  because  of 
the  damage  caused  to  munition  factories,  transportation 
lines,  store  houses,  etc.,  but  also  because  of  the  moral 
effect  on  both  the  enemy  troops  and  the  civil  population. 
A  well-timed  bombing  raid  will  do  more  to  disorganize 


MILITARY  AEROPLANES 


43 


44  MILITARY  AEROPLANES 

an  army  than  almost  any  other  form  of  attack,  and  this 
is  attended  with  a  much  less  loss  of  life,  and  with  less 
cost  and  equipment.  Points  in  enemy  territory  that  could 
be  reached  in  no  other  way  are  readily  attacked  by  bomb- 
ing planes  with  all  the  disastrous  effects  of  heavy  artillery 
fire.  The  aeroplane  is  better  adapted  for  this  service  than 
dirigibles  of  the  Zeppelin  type,  for  they  require  fewer  men 
for  their  operation,  and  in  addition  cost  less  to  operate 
and  build. 

Bombing  operations  against  well  protected  objectives 


Fig.   4a,     Curtiss   "Wireless"  Speed   Scouts    (S-2).     By  an   Ingenious  Arrange- 
ment of  the  Interplane  Struts  There  Is  No  Exposed  Wire  or  Cable. 

are  best  made  at  night  since  there  is  less  chance  of  loss 
through  anti-aircraft  gun  fire,  and  also  because  of  the 
difficulty  that  the  defense  machines  have  in  locating  the 
raiders.  Even  when  well  equipped  with  searchlights  and 
listening  stations,  it  is  not  the  easiest  thing  in  the  world 
to  pick  out  and  hold  the  location  of  an  attacking  squadron, 
for  the  searchlights  immediately  betray  them_selves  and 
can  then  be  put  out  of  action  by  fire  from  the  invaders. 
With  the  searchlights  out  of  commission,  it  is  almost 
impossible  for  the  defending  chasers  to  locate  and  engage 
the  raiders,  even  before  the  bombs  have  been  dropped. 
After  the  bomb  dropping  has  been  accomplished    (and 


46 


MILITARY  AEROPLANES 


with  comparative  accuracy  because  of  the  flares  dropped 
by  the  bombing  party),  the  raiders  are  lightened  of  a 
considerable  portion  of  their  load,  and  are  correspond- 
ingly increased  in  their  ability  to  climb  and  to  evade  the 
enemy  chasers. 

Night  flying  in  squadrons  always  introduces  the  danger 
of  collision,  and  to  minimize  this  danger,  by  decreasing 
the  number  of  machines,  the  size  and  carrying  capacity 
of  the  bombers  has  been  continually  increased.     Again, 


Fig.   6. 


Nieuport   Biplane    Scout  with   Machine  Gun   Pivoted  Above  the  Upper 
Wing.     This  Gun  Fires  Above  the  Propeller. 


bombing  requires  the  steady  platform  that  only  a  large 
machine  can  give,  and  for  accuracy  the  span  and  area 
must  be  greater  than  that  of  the  reconnaissance  type. 
In  night  flying  a  large  machine  is  safer  to  handle  owing 
to  its  lower  landing  speed  and  ability  to  come  to  rest 
quickly  after  landing,  and  this  is  of  the  greatest  im- 
portance when  landing  outside  of  the  aerodrome.  For 
daylight  work  at  comparatively  short  distances  the 
smaller  bomb  carrier  used  at  the  beginning  of  the  war 
is  probably  preferable  as  it  has  better  maneuvering 
qualities,  and  as  the  bombs  are  divided  among  a  greater 


MILITARY  AEROPLANES 


47 


number  of  machines  they  are  not  so  likely  to  be  de- 
feated before  accomplishing  their  object.  Because  of 
their  great  size,  these  bombing  aeroplanes  are  nearly 
always  of  the  "twin  motor"  type  with  two,  or  even  three, 
independent  power  plants.  The  use  of  a  twin  power 
plant  is  an  added  insurance  against  forced  landings  in 
hostile    country,   or   over   unsuitable   ground,   and   even 


Fig.  8.  Fokker  Synchronized  Machine  (Inn.  The  Gun  Is  Driven  by  the  Motor 
in  Such  a  Way  That  the  Bullets  Pass  Between  the  Propeller  Blades. 
"L'Aerophile." 

with  one  dead   engine  the  machine  can  be  flown   home 
at  a  fair  speed. 

"Spotting"  for  the  guidance  of  the  artillery  is  a  ciury 
usually  performed  by  the  reconnaissance  type,  or  small 
bombing  type,  and  is  usually  done  under  the  escort  of 
chasers.  Their  duty  is  to  direct  the  battery  as  to  the 
placing  of  shots.  The  ideal  machine  for  such  a  purpose 
would  be  the  direct  lift  type  similar  to  a  helicopter  which 
could  hover  over  one  particular  spot  until  its  object  had 
been  accomplished  in  making  measurements,  and  plot- 
ting enemy  positions.  Since  no  such  machine  is  at 
present  available,  the  duty  must  be  performed  by  a  low 
speed  aeroplane,  that  is  large  enough  to  provide  a  fairly 


48  MILITARY  AEROPLANES 

Steady  platform  and  at  the  same  time  has  sufficient  speed 
for  a  quick  getaway.  A  dirigible  has  the  necessary  hover- 
ing qualities  but  lacks  the  speed  necessary  for  avoiding 
attack  from  even  the  slowest  of  aeroplanes,  and  in  addi- 
tion is  a  magnificent  target  for  anti-craft  guns  if  kept  at 
an  altitude  low  enough  for  accurate  observation.  A 
large  speed  range  is  a  desirable  characteristic  in  such 
service. 

Photography  is  of  the  greatest  importance  in  recon- 
naissance, since  the  camera  distinctly  records  objects  on 
the  terrain,  so  small  and  obscure  that  they  may  entirely 
escape  the  eye  of  the  observer.  Again,  the  photograph 
is  a  permanent  record  that  may  be  studied  at  leisure 
in  headquarters,  or.  may  be  used  in  comparisons  with 
photographs  taken  at  an  earlier  date  in  the  same  terri- 
tory. Thus  changes  in  the  disposition  of  enemy  bat- 
teries, trenches,  and  troops  can  be  quickly  identified. 
With  modern  aeroplane  photographic  equipment,  a  vast 
territory  may  be  investigated  and  mapped  out  by  a  single 
machine  in  a  few  hours.  Camouflage  has  but  few  ter- 
rors for  the  camera,  and  the  photographs  often  clearly 
reveal  that  which  has  been  passed  over  time  and  time 
again  by  the  observers.  When  sent  out  on  a  specific 
mission,  the  aeroplane  returns  the  films  in  an  amazingly 
short  length  of  time,  and  within  a  few  minutes  they  are 
developed  and  are  ready  for  the  inspection  of  the  officers 
in  charge.  The  analysis  of  these  photographs  has  rapidly 
developed  into  a  science  well  worthy  of  a  Sherlock 
Holmes.  Changes  in  the  position  of  shadows,  suspiciously 
sudden  growths  of  underbrush,  changes  in  the  direction 
of  paths,  and  fresh  mounds  of  earth  all  have  a  definite 
meaning  to  the  photographic  expert. 

In  the  navy  the  aeroplane  has  proved  of  much  value 
in  scouting  and  particularly  in  defense  against  the  sub- 
marine. Because  of  its  great  speed  it  has  a  daily  radius 
of  action  many  times  that  of  a  torpedo  boat,  and  because 


MILITARY  AEROPLANES 


49 


Fig.  9.  Types  of  Aeroplane  Bombs.  The  Tail  Surfaces  Guide  the  Bomb  So 
That  It  Strikes  on  the  Firing  Pin  and  at  the  Same  Time  "Safeties" 
the  Bomb  So  That  it  Will  Not  Explode  Until  it  Has  Fallen  for  Some 
Distance.  In  Falling,  the  Tail  Blades  Rotate  and  Release  the  Firing 
Mechanism  After  the  Bomb  Has  Fallen  Clear  of  the  Aeroplane.  Cour- 
tesy of  "Flying." 


oO 


MILITARY  AEROPLANES 


of  its  altitude  the  effective  range  of  vision  is  still  further 
increased.  At  a  fair  height  the  observer  can  easily  detect 
a  submarine  even  w^hen  submerged  to  a  considerable 
depth,  a  feat  impossible  when  near  the  sea  level.  For 
disclosing  the  conditions  existing  in  an  enemy  harbor 
the  aeroplane  is  fully  the  equal  of  the  dirigible  since  it 
can  approach  and  retreat  rapidly,  and  v^athout  much 
danger  at  comparatively  low  altitudes.     While  the  diri- 


Fig.   9-a.      Curtiss    "JN"   Twin    Motor    Biplane.     Observer    Is    Seated    in    Front. 


gible  can  float  indefinitely  at  one  point,  it  must  be  done 
at  an  altitude  that  is  safely  out  of  range  of  the  enemy 
guns,  and  this  is  usually  at  a  point  where  observation  is 
a  difficult  proposition.  It  does  not  take  long  to  get  the 
range  of  such  a  target  as  a  hovering  dirigible,  yet  at 
a  much  lower  altitude  it  is  difficult  to  handle  naval  anti- 
aircraft guns  effectively  against  a  speeding  aeroplane. 
The  smaller  scouting  seaplanes  can  report  the  position 
of  a  submarine  to  a  torpedo  boat  or  *'sub-chaser,"  while 
the  larger  machines  are  perfectly  capable  of  dealing  with 


MILITARY  AEROPLANES 


51 


52  MILITARY  AEROPLANES 

the  submarine  at  first  hand.  On  the  large  bombing  type, 
a  three-pounder  gun  and  a  number  of  large  bombs  can 
be  carried,  either  of  which  would  be  sufficient  for  the 
purpose. 

In  land  defense  chasers  and  fighters  are  used  for  patrol, 
and  to  maintain  a  barrage  against  the  entrance  of  enemy 
machines  into  our  lines.  The  patrol  machines  work 
along  the  front  line  trenches,  while  the  machines  main- 
taining the  barrage  are  generally  arranged  in  two  parallel 
lines  back  of  the  trenches,  the  first  being  about  five  miles, 
and  the  second  about  ten  miles  from  the  front.  All  three 
lines  are  generally  placed  between  the  enemy  and  the 
principal  stations  and  railroad  centers  to  insure  protection 
from  enemy  bombers  and  reconnaissance  machines. 
Should  the  first  Hne  patrol  fail  to  keep  raiders  from 
crossing  the  first  line  trenches,  they  will  have  to  pass 
through  at  least  two  more  zones  of  organized  fighting 
squadrons  before  reaching  a  vulnerable  spot  in  our  lines. 
The  machines  used  for  patrol  and  barrage  are  of  the 
high  speed  and  fast  climbing  chaser  type.  The  response 
to  an  attack  involves  rapid  climbing,  and  a  high  degree 
of  maneuvering. 

Except  for  the  bombers  and  battle  planes,  the  machine 
gun  or  "Mitraleuse"  has  been  the  only  form  of  arm  in 
common  use  on  aeroplanes.  These  use  ammunition  ap- 
proximating service  rifle  caliber  and  are  furnished  in 
bands,  strips  or  drums  according  to  the  type  of  gun. 
With  larger  guns,  the  weight  of  the  ammunition  has 
been  found  excessive  with  all  but  the  largest  bombing 
machines,  and  the  recoil  of  a  large  caliber  gun  has  also 
been  difficult  to  overcome.  In  a  modern  American  aero- 
plane gun  of  large  caliber  the  recoil  has  been  reduced  to 
almost  a  negligible  degree,  even  up  to  the  four-pounder 
size,  by  a  system  of  balanced  projectile  reactions.  This 
gun  has  met  successful  tests,  but  whether  it  has  met  with 
general  adoption  would  be  difficult  to  say  at  the  present 


MILITARY  AEROPLANES 


53 


I^H 


•5^5 

3   «  g 

'Z  "<*i 

"C  >  o 
O-r  i- 


=    ">    r- 

^^-^ 

C  1-   O 


o  o  5< 

O         U 


c  2 


54  MILITARY  AEROPLANES 

time.  In  Europe,  large  caliber  aeroplane  guns  have  been 
used  on  large  "battle  planes"  or  ''gun  planes"  for  shelling 
dirigibles,  or  in  destroying  searchlight  stations  in  bomb- 
ing raids.  The  battle  planes  are  nearly  always  of  the 
**Twin"  type  with  the  gun  mounted  in  the  front  end  of 
the  fuselage. 

Summary  of  Types.  To  sum  up  the  types  required 
in  military  operations,  we  have:  (1)  High  speed 
''Chaser"  or  "Scout"  (Single  seater),  (2)  High  speed 
"Chaser"  (Two-seat  type),  (3)  Reconnaissance  type, 
(4)  Bombing  type,  (5)  Gun  or  Battle  Planes.  This  does 
not  include  the  training  machines  of  the  two  place  and 
"Penguin"  types,  but  as  these  are  simply  unarmed  modi- 
fications of  the  two  place  reconnaissance  and  single  seat 
machine,  respectively,  we  will  not  go  into  further  details 
at  this  point  regarding  their  construction. 

The  Chaser  or  Pursuit  Type.  The  most  important 
factors  in  the  design  of  a  chaser  are  speed  and  maneuver- 
ing ability.  The  speed  must  be  at  a  maximum  in  both 
the  horizontal  and  vertical  directions,  for  cHmbing  ability 
is  fully  of  as  much  importance  as  horizontal  speed.  Sec- 
ond in  importance  is  the  maximum  altitude  or  "Height 
of  ceiling"  to  which  the  machine  can  ascend.  This  maxi- 
mum "Ceiling"  generally  goes  hand  in  hand  with  the 
climbing  speed,  since  a  fast  climber  generally  has  a 
correspondingly  high  maximum  altitude.  The  combina- 
tion of  weight  and  head  resistance  must  be  such  that  the 
climb  interfers  as  little  as  possible  with  the  forward 
velocity. 

Great  climbing  ability  means  a  large  power  reserve, 
hence  the  weight  carried  per  horsepower  is  reduced  to 
from  8  to  12  pounds  on  the  fastest  machines,  against  the 
16  to  18  pounds  carried  on  the  larger  and  slower  recon- 
naissance types.  Strength  must  be  sacrificed  to  meet 
these  conditions,  so  that  instead  of  having  a  safety  factor 
of  from  8  to  12  as  in  the  larger  machines,  it  is  cut  down 


MILITARY  AEROPLANES 


55 


56  MILITARY  AEROPLANES 

to  about  5.5,  or  in  other  words,  the  strength  is  relatively 
only  half  that  of  the  usual  type  of  aeroplane.  This  great 
reduction  in  strength  calls  for  careful  handling,  especially 
in  landing,  and  also  painstaking  care  in  the  design  and 
choice  of  materials. 

High  speeds  and  maneuvering  ability  both  call  for 
small  wing  areas  and  short  spans,  the  areas  being  so 
adjusted  that  the  resistance  is  at  a  minimum  at  the  high- 
est speeds.  The  short  spans  have  a  minimum  of  exposed 
interplane  bracing  and  thus  indirectly  reduce  both  the 
head  resistance  and  the  weight.  Unfortunately,  the  most 
favorable  areas  at  high  speeds  are  too  small  for  safe 
landing  speeds.  With  a  fixed  area,  the  minimum  landing 
speed  is  only  a  little  less  than  one-half  of  the  maximum 
flying  speed,  hence  with  a  maximum  of  150  miles  per 
hour  the  minimum  will  probably  be  little  less  than  70 
miles  per  hour.  The  most  efficient  wing  sections,  and 
the  greatest  refinement  in  the  body  design,  bracing,  and 
chassis  are  necessary  at  speeds  of  over  100  miles  per 
hour.  All  other  conditions  being  equal,  the  resistance 
varies  as  the  square  of  the  velocity,  hence  at  150  miles 
per  hour,  the  resistance  is  2.25  times  that  at  100  miles 
per  hour. 

The  following  table  gives  approximately  the  appor- 
tionment of  the  head  resistance  producing  items  in  a 
typical  speed  scout  or  chaser. 

Body  (Fuselage)   68  per  cent 

Chassis,  wheels,  struts,  etc 15  per  cent 

Tail,  rudder,  fin,  elevator 5  per  cent 

Wing  structure,  struts,  wire,  fittings .  12  per  cent 

The  aerodynamic  drag  due  to  the  lift  of  the  wings  is 
not  included  in  the  above,  the  useless  or  parasitic  resist- 
ance alone  being  considered.  It  will  be  noted  that  the 
body  causes  by  far  the  greater  part  of  the  resistance,  and 
as  a  result,  the  body  of  the  speed  scout  requires  the  most 


MILITARY  AEROPLANES  57 

careful  attention  in  regard  to  streamline  form.  Fortun- 
ately this  is  possible  with  the  short  stumpy  body  of  the 
chaser,  since  a  true  streamline  form  approximates  the 
average  body  dimensions  of  the  scout.  In  the  larger 
machines,  the  body  resistance  is  not  as  great  in  propor- 
tion to  the  other  items  since  there  are  more  struts  and 
stay  wires,  the  chassis  is  larger,  and  the  tail  surfaces  are 
of  greater  area.  The  chassis  is  the  next  largest  item  and 
is  one  of  the  most  difficult  items  to  reduce.  It  has  been 
suggested  by  several  people  that  the  chasis  could  be  stored 
away  in  the  body  while  in  flght,  but  this  adds  additional 
mechanism  and  weight,  and  any  automatic  mechanism  for 
folding  up  the  chassis  members  would  likely  prove  un- 
reliable. 

Chaser  Armament.  A  single  seat  chaser  is  provided 
with  one  or  two  machine  guns  mounted  on  top  of  the 
fuselage,  and  directly  in  front  of  the  pilot,  the  length  of 
the  barrel  being  parallel  with  the  fore  and  aft  center  line. 
They  may  either  be  fixed  rigidly  to  the  fuselage  top,  or  so 
that  they  can  be  pointed  up,  and  over  the  top  of  the  upper 
wing.  With  the  machine  guns  fixed  rigidly  to  the  body, 
as  in  the  early  chaser  monoplanes  used  by  Garros  and 
Vedrines,  it  was  necessary  at  all  times  to  fire  directly 
through  the  disc  area  swept  out  by  the  propeller. 

Two  plans  were  tried  for  preventing  the  propeller  from 
being  broken  by  the  bullets.  The  first  consisted  of  a 
device  operated  by  the  motor  that  stopped  the  gun  when- 
ever the  propeller  blade  came  within  the  path  of  the  bul- 
lets. This  early  mechanism  proved  unreliable,  since  the 
frequent  stopping,  with  the  propellers  running  1200  revo- 
lutions per  minute,  soon  put  the  apparatus  out  of  order. 
Soon  after  the  failure  of  this  method,  designers  mounted 
curved  protective  steel  plates  on  the  inner  portions  of  the 
propeller  blades  at  points  where  they  were  likely  to  be 
struck  with  bullets.  According  to  calculations  in  proba- 
bility and  chance,  only  one  bullet  out  of  every  eighteen 


58  MILITARY  AEROPLANES 

will  strike  the  protective  plate  on  the  propeller  blade,  and 
hence  only  one  out  of  eighteen  bullets  will  be  wasted. 
This,  however,  was  a  makeshift,  and  on  modern  machines 
the  gun  is  driven,  or  ''Synchronized"  with  the  motor  so 
that  the  bullets  pass  between  the  blades. 

Many  modern  single  seat  chasers  have  the  gun  pivoted 
to  the  top  of  the  fuselage  so  that  the  pilot  can  fire  above 


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Curtiss  Biplane  in  Flight.    Taken  from  Another  Machine.    Courtesy  "Aerial  Age." 

the  top  plane  and  to  either  side  of  the  body.  This  does 
away  with  the  difficulty  of  keeping  the  machine  headed 
directly  at  the  enemy  when  in  action,  a  method  that  is 
imperative  with  the  fixed  type  of  gun.  Two  seater  chasers 
are  generally  arranged  so  that  the  gunner  is  seated  back 
of  the  pilot,  and  the  gun  is  so  pivoted  and  supported  that 
it  can  be  swung  through  a  wide  radius  both  toward  the 
front  and  on  either  side.  This  freedom  of  gun  action  ;.t 
least  partly  compensates  for  the  slower  maneuvering 
qualities  of  the  two  seater  type,  since  the  gun  may  be 
swung  with  the  target  through  quite  a  range  of  field,  and 


MILITARY  AEROPLANES  59 

without  changing  the  flight  direction  of  the  machine.  A 
gun  of  this  type  is  provided  with  stops  which  prevent  the 
gunner  from  shooting  into  the  outlying  parts  of  his  own 
machine.  The  gun  mounting  in  many  cases  of  two  seater 
construction  consists  of  a  light  circular  track  that  runs 
around  the  edge  of  the  cockpit  opening.  The  gun  stand- 
ard runs  on  this  track,  and  the  gun  is  pivoted  at  the  upper 
end  of  the  standard  so  that  the  muzzle  can  be  raised  or 
lowered.  The  gun  turns  in  a  horizontal  plane  by  sliding 
on  the  track,  and  can  be  followed  around  by  the  gunner 
who  is  seated  in  the  center  on  a  pivoted  seat.  With  this 
mounting  it  is  possible  to  guard  against  a  rear  attack,  to 
shoot  straight  up.,  or  nearly  straight  down  over  the  sides 
of  the  fuselage. 

In  a  few  machines  of  the  two  seater  type,  two  machine 
guns  are  provided,  one  pivoted  gun  in  the  rear,  and  one 
gun  rigidly  fastened  to  the  fuselage  in  front  of  the  pilot. 
It  is  very  seldom  that  both  guns  can  be  brought  into 
action  at  once  unless  engaged  with  a  number  of  enemy 
machines,  although  the  front  gun  is  handy  in  pursuit,  and 
at  a  time  when  the  rear  gun  is  ineffective  because  of  the 
pilot  in  front  of  him.  Even  with  the  double  equipment, 
the  superior  maneu\ering  qualities  of  the  single  seater 
makes  matters  more  even  than  would  commonly  be  sup- 
posed. An  added  advantage  of  the  single  seater  is  that 
it  is  smaller  and  therefore  more  difficult  to  hit. 

English  speed  scouts  have  largely  adopted  the  Amer- 
ican Lewis  gun.  The  cartridges  in  this  gun  are  arranged 
radially  in  a  circular  drum,  and  are  fed  to  the  gun  as  the 
drum  revolves.  The  drum  is  mounted  on  the  barrel  near 
the  breech  and  is  operated  automatically  by  the  successive 
explosions.  This  feeds  the  cartridges  and  rejects  the 
empty  shells  without  the  attention  of  the  pilot.  It  fires 
about  600  shots  per  minute.  When  one  drum  is  ex- 
hausted, another  drum  of  new  cartridges  can  be  quickly 
and  easily  inserted.     The  French  use  the  belt  system  to 


60  MILITARY  AEROPLANES 

a  large  extent.  In  this  system  the  cartridges  are  attached 
side  by  side  on  a  cotton  web  belt  as  in  the  older  types  of 
army  machine  guns.  As  in  the  Lewis  gun,  the  cartridges 
are  fed  automatically  by  the  recoil  of  the  explosions,  and 
the  belt  moves  through  the  breech  with  a  step  by  step 
movement  until  the  ammunition  is  exhausted.  This  is  not 
nearly  as  compact  an  arrangement  as  the  Lewis  gun,  and 
is  more  difficult  to  pivot  on  account  of  the  dangling  belts. 

On  the  right  hand  side  of  the  Nieuport  body  there  is  a 
drum  on  which  the  belt  with  the  loaded  cartridges  is 
wound.  The  empty  end  of  the  belt  is  wound  on  a  drum 
at  the  left,  this  drum  being  provided  with  a  spring  to  keep 
the  belt  taut.  The  empty  cartridges  are  discharged 
through  a  tube  that  passes  through  the  side  of  the  body. 
On  the  1916  Fokker  the  gun  is  of  the  Maxim  type,  and  is 
immovably  fastened  above  the  engine  cowl  and  slightly 
to  the  right.  To  fire  the  gun,  the  pilot  presses  down  a 
small  lever  fastened  to  the  control  column,  and  from  this 
lever  the  connecting  Bowden  wire  closes  the  motor  clutch 
and  starts  the  gun.  A  cam  is  fixed  to  the  motor  shaft  in 
relation  to  the  propeller  blades.  When  firing,  the  elevator 
control  is  locked  fore  and  aft,  while  the  lateral  control 
movement  is  operated  by  the  pilot's  knees.  Steering  is  by 
the  action  of  his  feet  on  the  rudder  bar.  Thus  the  pilot 
can  balance  laterally,  and  steer  with  his  hands  free  for  the 
manipulation  of  the  gun,  but  he  cannot  change  his  eleva- 
tion. 

Power  Plant  of  the  Chaser.  In  the  smaller  speed  scouts, 
the  motor  is  of  the  rotary  air  cooled  type,  the  output 
ranging  from  80  to  110  horsepower,  but  as  the  power 
demands  increased  the  water-cooled  motor  came  into  use, 
and  at  the  present  time  has  found  favor  with  a  large  num- 
ber of  builders.  When  the  power  exceeds  about  120  horse- 
power it  is  difficult  to  thoroughly  cool  the  rotary  engine, 
and  although  the  Gnome,  Clerget  and  Le  Rhone  are  ex- 
tremely desirable  on  a  chaser  because  of  their  light  weight, 


MILITARY  AEROPLANES 


61 


62  MILITARY  AEROPLANES 

they  cannot  be  used  profitably  on  the  larger  scouts.  Up 
to  the  present  time,  the  Nieuport  and  Sopwith  use  the 
Clerget  and  Le  Rhone  rotary  motors,  but  the  S.  P.  A.  D. 
and  several  others  have  adopted  the  water-cooled  type. 
Nearly  all  of  the  German  chasers,  such  as  the  Roland  and 
Albatros,  are  v^ater-cooled.  Such  motors  must  weigh  well 
below  3  pounds  per  horsepower  if  there  is  to  be  sufficient 
power  reserve  for  fast  climbing.  The  Curtiss  scouts  are 
also  water-cooled,  although  the  rating  is  only  100  horse- 
power. The  French  and  German  machines  are  very 
heavily  powered,  motors  of  175  horsepower  being  very 
common,  even  on  single  seaters.  The  fuel  capacity  is  very 
limited,  probably  not  exceeding  2.5  to  3  hours  in  any  case. 

General  Dimensions  of  Scouts.  The  following  table 
will  give  a  better  idea  of  the  principal  characteristics  of 
these  machines.  It  gives  the  overall  dimensions,  power, 
speed,  climb,  etc.  It  will  be  noted  that  the  Nieuport  bi- 
plane scouts  have  a  smaller  lower  chord  (*).  The  speeds 
given  are  the  sea  level  speeds  since  a  great  change  in  alti- 
tude afifects  the  performance  to  a  marked  degree. 

Reconnaissance  Type  Arrangement.  These  machines 
are  almost  invariably  of  the  two  seater  type,  and  are 
equipped  either  with  one  machine  gun  for  the  observer,  or 
with  a  rigidly  fixed  gun  for  the  pilot  and  a  pivoted  gun  at 
the  rear  for  the  observer.  In  the  majority  of  cases  the 
observer  is  seated  in  the  rear  cockpit  (Tractor  types), 
and  at  a  point  where  he  has  a  greater  visual  radius  and 
field  of  fire.  With  the  pusher  type,  the  observer  is,  of 
course,  seated  in  the  extreme  front  of  the  body,  where  he 
has  an  extremely  wide  angle  of  vision.  The  pilot  in  the  rear 
seat  of  the  pusher  is  effectually  screened  from  any  gun 
action,  either  from  the  front,  side  or  rear,  as  the  propeller 
cuts  off  the  field  at  the  back  and  the  observer  and  inter- 
plane  bracing  blocks  the  way  at  the  front  and  sides.  The 
observer's  cock-pit  is  equipped  with  the  signalling  appa- 
ratus, photographic  equipment,  map  boards,  etc.,  as  well 


MILITARY  AEROPLANES 


63 


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64  MILITARY  AEROPLANES 

as  the  ammunition  for  the  gun.    The  pilot's  compartment 
contains  the  navigating  instruments  and  controls. 

Armament.  At  the  beginning  of  the  war  nearly  all  of 
the  French  two  seaters  were  of  the  pusher  type,  this 
arrangement,  of  course,  resulted  in  almost  a  completely 
dead  angle  of  fire  in  the  rear,  and  a  front  horizontal  angle 
that  was  practically  restricted  to  160  degrees.  Owing  to 
the  forward  position  of  the  gun  the  vertical  angle  was 
quite  good,  230  degrees  or  even  better.  In  the  tractor 
two  seater,  with  a  single  movable  gun  mounted  "En  bar- 
bette" at  the  rear,  the  horizontal  angle  is  about  180  de- 
grees, but  the  vertical  angle  is  less  than  with  the  pusher 
type.  When  the  rear  gun  is  supplemented  with  a  front 
rigidly  mounted  gun,  there  is  some  protection  at  the  front, 
but  the  rigid  gun  is  far  from  being  as  effective  as  the 
pivoted  rear  gun.  The  front  gun  of  course  fires  through 
the  propeller.  This  armament  is  used  by  the  German 
machines  "Aviatic,"  "Rumpler,"  "Albatros,"  and  "L.  V. 
G."  The  forward  rigid  gun  is  usually  of  the  infantry 
type,  while  the  movable  rear  gun  is  lighter.  The  latter  is 
fed  by  drums,  or  rolled  bands  on  spools,  so  that  reloading 
can  be  performed  in  the  wind  stream. 

With  the  two  seater  type  used  in  reconnaissance,  artil- 
lery spotting,  or  photography,  the  power  is  generally  in 
the  neighborhood  of  220-260  horsepower,  and  the  speed 
varies  from  85  to  100  miles  per  hour.  The  area  is  approxi- 
mately 400  to  480  square  feet.  A  single  engine  is  gen- 
erally used. 

General  Dimensions  and  Speeds.  Reconnaissance  ma- 
chines of  various  types  and  makes  are  listed  in  the  follow- 
ing table.  A  pusher  is  indicated  by  (P)  and  a  tractor  by 
(T).  The  German  aeroplanes  (G),  and  the  Allied  aero- 
planes (A),  are  both  listed  for  comparison: 

It  will  be  noted  that  several  types  of  machines  have 
been  made  by  the  same  firms,  and  that  in  some  cases  the 
same  machines  have  different  power  plants.     The  Alba- 


MILITARY  AEROPLANES 


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66  MILITARY  AEROPLANES 

tros  C-III  has  been  furnished  with  both  the  170  and  220 
Mercedes  motor.  The  Ago  biplane  has  a  tapering  wing, 
and  the  chord  width  (*)  given  is  taken  at  the  body.  While 
very  recent  machines  cannot  be  described,  because  of  cer- 
tain restrictions,  the  horsepower  of  the  latest  two  seaters 
will  average  about  240  horsepower.  If  the  dates  and 
power  items  are  noted,  it  will  be  seen  that  the  machines 
used  in  1917  have  much  larger  motors  than  those  built  in 
1916.  The  weight  per  square  foot  of  surface  will  average 
about  6.5  pounds.  The  loading  per  horsepower  rarely 
exceeds   17.0  pounds. 

Bombing  Type  Aeroplanes.  These  large  aeroplanes  are 
fitted  with  either  two  or  three  independent  power  plants. 
The  German  bombers  are  represented  by  the  Gotha, 
A.  E.  G.,  Friedrichshafen,  and  Rumpler  G,  while  the  Allied 
bombers  are  the  Caproni,  Handley-Page,  Farman,  Voisin, 
etc.  The  speed  is  about  that  of  the  reconnaissance  type, 
and  will  seat  three  or  more  men.  The  motors  average  500- 
560  horsepower  per  power  plant,  and  the  wing  area  is 
usually  well  over  1,000  square  feet.  The  small  two  seaters 
are  generally  equipped  with  two  pivoted  machine  guns, 
while  the  three  seaters  have  a  third  machine  gun  ar- 
ranged so  that  it  can  be  lowered  and  fired  through  a  trap 
door  in  the  bottom  of  the  body.  Defense  may  thus  be  had 
from  the  rear,  or  below.  In  some  of  the  pusher  types,  a 
rapid  fire  gun  of  comparatively  heavy  caliber  is  mounted 
at  the  front  of  the  body  in  place  of  the  usual  machine  gun. 
This  is  usually  the  case  with  the  sea  planes  used  for  sub- 
marine chasing. 

In  addition  to  bombing  operations,  these  large  ma- 
chines are  also  used  for  the  protection  of  ''spotting"  aero- 
planes, or  for  the  direct  protection  of  the  lines  against 
land  attacks.  These  heavily  armed  bombers  are  very  diffi- 
cult to  attack,  even  for  the  smaller  and  more  agile  "Chas- 
ers," as  they  can  fire  from  below  as  well  as  from  the 
front,  top,  or  sides.     In  the  bombers  which  have  only  a 


MILITARY  AEROPLANES  67 

single  gun  in  the  rear,  the  gunner  is  working  at  a  disad- 
vantage if  his  adversary  forces  him  to  continually  raise 
and  lower  his  gun  from  the  top  of  the  body  to  the  lower 
trap  door.  This  is  very  tiring  to  the  rear  gunner,  and  if 
the  chaser^s  tactics  are  carried  out  for  a  sufficient  length 
of  time,  it  can  wear  out  the  gunner  by  continually  rising 
and  dropping  at  the  tail  of  the  bombing  plane.  In  regard 
to  the  front  gun,  the  twin  motor  type  ofifers  many  of  the 
advantages  of  the  pusher,  and  as  a  whole,  the  twin  ar- 
rangement will  nearly  double  the  field  of  fire  of  either 
the  tractor  or  pusher. 

The  bombing  planes  must  have  a  very  large  radius  of 
action,  particularly  those  that  are  used  in  night  bombing 
operations.  The  Gothas  in  bombing  London  fly  several 
hundred  miles  from  their  base,  and  recently  a  Handley- 
Page  bombing  plane  flew  from  London  to  Constantinople, 
Turkey,  making  only  a  few  stops  on  the  way.  Starting 
out  from  Hendon,  England,  the  Handley-Page  machine 
flew  to  Paris,  down  the  Rhone  valley  to  Lyons  and  ]Mar- 
seilles,  and  then  to  Pisa  and  Rome  (Italy),  where  they 
landed.  From  Rome  the  machine  passed  over  Naples, 
over  Oranto  and  then  over  the  Albanian  Alps  to  the 
base  at  Salonica.  flaking  preparations  at  this  base  they 
flew  the  final  stage  of  the  trip  to  Constantinople,  a  dis- 
tance of  250  miles  over  hostile  country.  The  bombing  of 
the  Turkish  capital  was  done  at  night  after  a  flight  of 
21/2  hours  from  Salonica.  When  over  the  sea  of  Marmora, 
the  ship  "Goeben"  was  bombed,  and  in  addition  a  hit  was 
scored  on  the  two  submarines  lying  at  her  side.  Four 
bombs  struck  the  **Goeben"  directly,  from  an  altitude  of 
800  feet.  Two  more  bombs  were  dropped  on  the  German 
ship,  ''General,"  which  was  the  headquarters  of  the  Ger- 
man staff.  Finally,  after  30  minutes  over  the  city  of  Con- 
stantinople, the  Turkish  War  Office  was  the  recipient  of 
two  more  bombs.  In  the  words  of  the  Turkish  com- 
munique this  ''Was  not  entirely  destroyed."     On  its  re- 


68  MILITARY  AEROPLANES 

turn  to  Salonica  it  was  found  that  the  machine  had  been 
struck  by  26  shrapnel  bullets.  This  disabled  one  of  the 
power  plants  so  that  the  greater  part  of  the  return  journey 
was  made  on  a  single  m^tor. 

From  London  to  Salonica  five  men  were  carried.  In 
addition  was  their  luggage,  bedding,  two  tool  boxes,  spare 
parts  equivalent  in  weight  to  one  engine,  and  two  11 '-6' 
spare  propellers.  Complete,  the  machine  weighed  over  6 
tons,  with  a  useful  load  of  about  6,000  pounds.  In  cross- 
ing the  Albanian  Alps  the  machine  frequently  was  at  an 
altitude  of  10,000  feet.  The  power  plant  consisted  of  two 
275  horsepower  Rolls-Royce  motors,  and  even  at  this  high 
altitude,  and  with  the  heavy  loading,  no  trouble  was  ex- 
perienced. During  the  bombing,  only  three  men  were 
carried,  the  remainder  of  the  useful  weight  being  made 
up  of  bombs  and  other  ammunition.  While  this  record 
will  probably  be  beaten  before  this  book  goes  to  press, 
it  will  at  least  give  an  idea  as  to  the  requirements  and 
capabilities  of  the  bombing  type  aeroplane. 

Military  Training  Machines.  The  military  training 
machines  used  in  the  United  States  are  generally  of  the 
two  seater  tractor  type,  similar  in  external  appearance  to 
the  reconnaissance  type  machines  already  described. 
They  are  low  powered,  90  to  125  horsepower,  and  will 
have  an  average  span  of  40'-0".  The  controls  are  in  dupli- 
cate so  that  the  student's  controls  move  in  unison  with 
the  instructor's. 


MILITARY  AEROPLANES 


69 


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70 


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CHAPTER  III. 
ELEMENTARY  AERODYNAMICS. 

Definition.  Aerodynamics  treats  of  the  forces  pro- 
duced by  air  in  motion,  and  is  the  basic  subject  in  the 
study  of  the  aeroplane.  It  is  the  purpose  of  this  chapter 
to  describe  in  detail  the  action  of  the  wing  in  flight,  and 
the  aerodynamic  behavior  of  the  other  bodies  that  enter 
into  the  construction  of  the  aeroplane.  At  present,  aero- 
dynamic data  is  almost  entirely  based  on  experimental 
investigations.  The  motions  and  reactions  produced  by 
disturbed  air  are  so  complex  and  involved  that  no  com- 
plete mathematical  theory  has  yet  been  advanced  that 
permits  of  direct  calculation. 

Properties  of  Air.  Air  being  a  material  substance,  pos- 
sesses the  properties  of  volume,  weight,  viscosity  and 
compressibility.  It  is  a  mechanical  mixture  of  the  two 
elementary  gases,  oxygen  and  nitrogen,  in  the  propor- 
tion of  23  per  cent  of  oxygen  to  77  per  cent  of  nitrogen. 
It  is  the  oxygen  element  that  produces  combustion,  while 
the  nitrogen  is  inert  and  does  not  readily  enter  into  com- 
bination with  other  elements,  its  evident  function  being  to 
act  as  a  dilutant  for  the  energetic  oxygen.  In  combus- 
tion, the  oxygen  enters  into  a  chemical  combination  with 
the  fuel  while  the  nitrogen  passes  off  with  the  products 
of  combustion  unchanged. 

Air  is  considered  as  a  fluid  since  it  is  capable  of  flowing 
like  water,  but  unlike  w^ater,  it  is  highly  compressible. 
Owing  to  the  difference  between  air  and  water  in  regard 
to  compressibility,  they  do  not  follow  exactly  the  same 
laws,  but  at  ordinary  flight  speeds  and  in  the  open  air, 
the  variations  in  the  pressure  are  so  slight  as  to  cause 

71 


72        ELEMENTARY  AERODYNAMICS 

little  difference  in  the  density.  Hence  for  flight  alone,  air 
may  be  considered  as  incompressible.  It  should  be  noted 
that  a  compressible  fluid  is  changed  in  density  by  varia- 
tions in  the  pressure,  that  is,  by  applying  pressure  the 
weight  of  a  cubic  foot  of  a  compressible  fluid  is  greater 
than  the  same  fluid  under  a  lighter  pressure.  This  is  an 
important  consideration  since  the  density  of  the  air 
greatly  affects  the  forces  that  set  it  in  motion,  and  for 
this  reason  the  density  (weight  per  cubic  foot)  is  always 
specified  in  a  test. 

Every  existing  fluid  resists  the  motion  of  a  body,  the 
opposition  to  the  motion  being  commonly  known  as 
"resistance."  This  is  due  to  the  cohesion  between  the 
fluid  particles  and  the  resistance  is  the  actual  force 
required  to  break  them  apart  and  make  room  for  the 
moving  body.  Fluids  exhibiting  resistance  are  said  to 
have  "viscosity."  In  early  aerodynamic  researches,  and  in 
the  study  of  hydrodynamics,  the  mathematical  theory  is 
based  on  a  "perfect  fluid,"  that  is,  on  a  theoretical  fluid 
possessing  no  viscosity,  and  while  this  conception  is 
an  aid  in  studying  the  reactions,  the  actual  laboratory 
results  are  far  from  the  computed  values.  Such  theory 
would  assume  that  a  body  could  move  in  a  fluid  without 
encountering  resistance,  which  in  practice  is,  of  course, 
impossible. 

In  regard  to  viscosity,  it  may  be  noted  that  air  is  highly 
viscuous — relatively  much  higher  than  water.  Density 
for  density,  the  viscosity  of  air  is  about  14  times  that  of 
water,  and  consequently  the  effects  of  viscosity  in  air  are 
of  the  utmost  importance  in  the  calculation  of  resistance 
of  moving  parts. 

Atmospheric  air  at  sea  level  is  about  1/800  of  the 
density  of  water.  Its  density  varies  with  the  altitude  and 
with  various  atmospheric  conditions,  and  for  this  reason 
the  density  is  usually  specified  "at  sea  level"  as  this  alti- 
tude gives  a  constant  base  of  measurement  for  all  parts  of 


ELEMENTARY  AERODYNAMICS        73 

the  world.  As  the  density  is  also  affected  by  changes  in 
temperature,  a  standard  temperature  is  also  specified. 
Experimental  results,  whatever  the  pressure  and  tempera- 
ture at  which  they  were  made,  are  reduced  to  the  corre- 
sponding values  at  standard  temperature  and  at  the  normal 
sea  level  pressure,  in  order  that  these  results  may  be 
readily  comparable  with  other  data.  The  normal  (aver- 
age) pressure  at  sea  level  is  14.7  pounds  per  square  inch, 
or  2,119  pounds  per  square  foot  at  a  temperature  of  60° 
Fahrenheit.  At  this  temperature  1  pound  of  air  occupies 
a  volume  of  13.141  cubic  feet,  while  at  0°  F.  the  volume 
shrinks  to  11.58  cubic  feet,  the  coresponding  densities 
being  0.07610  and  0.08633  pounds  per  cubic  foot,  respec- 
tively. This  refers  to  dry  air  only  as  the  presence  of  water 
vapor  makes  a  change  in  the  density.  With  a  reduc- 
tion in  temperature  the  pressure  increases  with  the  density 
increase  so  that  the  effect  of  heat  is  twofold  in  its  eft'ect. 

With  a  constant  temperature,  the  pressure  and  density 
both  decrease  as  the  altitude  increases,  a  density  at  sea 
level  of  0.07610  pounds  per  cubic  foot  is  reduced  to 
0.0357  pounds  per  cubic  foot  at  an  altitude  of  20,000  feet. 
During  this  increase  in  altitude,  the  pressure  drops  from 
14.7  pounds  per  square  inch  to  6.87  pounds  per  square  inch. 
This  variation,  of  course,  greatly  affects  the  performance 
of  aeroplanes  flying  at  different  altitudes,  and  still  more 
affects  the  performance  of  the  motor,  since  the  latter  can- 
not take  in  as  much  fuel  per  stroke  at  high  altitudes  as 
at  low,  and  as  a  result  the  power  is  diminished  as  we 
gain  in  altitude.  The  following  table  gives  the  power 
variations  at  different  heights  above  sea  level. 

This  air  table  also  gives  the  properties  of  air  through 
the  usual  range  of  flight  altitudes.  The  pressures  cor- 
responding to  the  altitudes  are  given  both  in  pounds  per 
square  inch  and  m  inches  of  mercury  so  that  barometer 
and  pressure  readings  can  be  compared.  In  the  fourth 
column  is  the  percentage  of  the  horsepower  available  at 


74 


ELEMENTARY  AERODYNAMICS 


different  altitudes,  the  horsepower  at  sea  level  being  taken 
as  unity.  For  example,  if  an  engine  develops  100  horse- 
power at  sea  level,  it  will  develop  100  X  0.66  =  66  horse- 
power at  an  altitude  of  10,000  feet  above  sea  level.  The 
barometric  pressure  in  pounds  per  square  inch  can  be 


TABLE  OF  ALTITUDES,  DENSITIES  AND  PRESSURES 

AT  60°  F. 


BAROMETRIC 

ALTITUDE  IN 

DENSITY  IN  LBS. 

PRESSURE     PER 

PERCENT       OF 

PRESSURE, 

FEET    ABOVE 

PER    CUBIC 

SQUARE 

POWER    (SEA 

IN  INCHES 

SEA      LEVEL 

FOOT 

INCH 
IN  POUNDS 

LEVBL= 
1.00 

OF   MER- 
CURY 

Sea 

Level^O.O 

0.07610 

14.701 

1.000 

30.00 

500 

0.07480 

14.333 

0.965 

29.25 

872 

0.07400* 

14.220 

0.950 

29.00 

1,000 

0.07340 

14.088 

0.945 

28.75 

1,340 

0.07250* 

13.965 

0.925 

28.50 

1,802 

0.07120* 

13.728 

0.913 

28.00 

2,000 

0.07070 

13.578 

0.903 

27.75 

2,753 

0.06900* 

13.238 

0.885 

27.00 

3,241 

0.06780* 

12.985 

0.878 

26.50 

4,276 

0.06540* 

12.495 

0.835 

25.50 

5,000 

0.06320 

11.956 

0.803 

24.40 

7,516 

0.05780* 

11.025 

0.748 

22.50 

10,000 

0.05230 

10.045 

0.660 

20.50 

12,630 

0.04740* 

9.165 

0.600 

18.50 

14,839 

0.04300* 

8.333 

0.550 

17.00 

18,109 

0.03810* 

7.350 

0.483 

15.00 

20,000 

0.03570 

6.86 

0.448 

14.00 

NOTE. — Densities  marked  *  are  interpolated  from  a  graph, 
but  are  close  enough  for  all  ordinary  purposes. 

obtained  by  multiplying  the  pressure  in  inches  of  mer- 
cury by  the  factor  0.4905,  this  being  the  v^eight  of  a  mer- 
cury column  1  inch  high. 

In  aerodynamic  laboratory  reports,  the  standard  density 
of  air  is  0.07608  pounds  per  cubic  foot  at  sea  level,  the 
temperature  being  15  degrees  Centrigrade  (59  degrees 
Fahrenheit).  This  standard  density  will  be  assumed 
throughout  the  book,  and  hence  for  any  other  altitude  or 
density  the  corresponding  corrections  must  be  made. 
Owing  to  the  fact  that  the  temperature  decreases  as  we 


ELEMENTARY  AERODYNAMICS 


iO 


gain  altitude,  further  corrections  must  be  made  in  the 
tabular  values,  but  as  the  changes  are  rather  difficult  to 
make  and  are  relatively  small  we  will  not  take  the  matter 
up  at  this  point. 

Air  Pressure  on  Normal  Flat  Plates.  When  a  flat  plate 
or  "plane"  is  held  at  right  angles  or  ''normal"  to  an  air 
stream,  it  obstructs  the  flow  and  a  force  is  produced  that 
tends  to  move  it  with  the  stream.     The  stream  divides. 


Fig.  1.  Air  Flow  About  a  Flat  Normal  Plate.  Pressure  Zone  at  Front  and 
Vacuous  Turbulent  Zone  at  Rear  (H).  Arrows  Show  Direction 
of  Flow. 

as  shown  in  Fig.  1  and  passes  all  around  the  edges  of 
the  plate  (P-R),  the  stream  reuniting  at  a  point  (M)  far 
in  the  rear.  Assuming  the  air  flow  from  left  to  right,  as 
in  the  figure,  it  will  be  noted  that  the  rear  of  the  plate 
at  (H)  is  under  a  slight  vacuum,  and  that  it  is  filled  with 
a  complicated  whirling  mass  of  air.  The  general  trend 
of  the  eddy  paths  are  indicated  by  the  arrows.  At  the 
front  where  the  air  current  first  strikes  the  plate  there  is 
a  considerable  pressure  due  to  the  impact  of  the  air  par- 
ticles. In  the  figure,  pressure  above  the  atmospheric  is 
indicated  by  *****^  while  the  vacuous  space  at  the  rear 
is  indicated  by  fine  dots.  As  the  pressure  in  front,  and 
the  vacuum  in  the  rear,  both  tend  to  move  the  surface  to 
the  right  in  the  direction  of  the  air  stream,  the  total  force 


76        ELEMENTARY  AERODYNAMICS 

tending  to  move  the  plate  will  be  the  difference  of  pres- 
sure on  the  front  and  rear  faces  multiplied  by  the  area  of 
the  plate.  Thus  if  F  is  the  force  due  to  the  impact  pres- 
sure at  the  front,  and  G  is  the  force  due  to  the  vacuum 
at  the  rear,  then  the  total  resistance  (D)  or  "Drag"  is  the 
sum  of  the  two  forces. 

Contrary  to  the  common  opinion,  the  vacuous  part  of 
the  drag  is  by  far  the  greater,  say  in  the  neighborhood  of 
from  60  to  75  per  cent  of  the  total.  When  a  body  expe- 
riences pressure  due  to  the  breaking  up  of  an  air  stream, 
as  in  the  present  case,  the  pressure  is  said  to  be  due  to 
^'turbulence,"  and  the  body  is  said  to  produce  "turbulent 
flow."  This  is  to  distinguish  the  forces  due  to  impact 
and  suction,  from  the  forces  due  to  the  frictional  drag 
produced  by  the  air  stream  rubbing  over  the  surface. 

Forces  due  to  turbulent  flow  do  not  vary  directly  as  the 
velocity  of  the  air  past  the  plate,  but  at  a  much  higher 
rate.  If  the  velocity  is  doubled,  the  plate  not  only  meets 
with  twice  the  volume  of  air,  but  it  also  meets  it  twice 
as  fast.  The  total  effect  is  four  times  as  great  as  in  the 
first  place.  The  forces  due  to  turbulent  flow  therefore 
vary  as  the  square  of  the  velocity,  and  the  pressure 
increases  very  rapidly  with  a  small  increase  in  the 
velocity.  The  force  exerted  on  a  plate  also  increases 
directly  with  the  area,  and  to  a  lesser  extent  the  drag  is 
also  affected  by  the  shape  and  proportions.  Expressed  as 
a  formula,  the  total  resistance  (D)  becomes:  D  =  KAV^, 
where  K^  co-efficient  of  resistance  determined  by  expe- 
riment, A=  area  of  plate  in  square  feet,  and  V=  velocity^ 
in  miles  per  hour.  The  value  of  K  takes  the  shape  and 
proportion  of  the  plate  into  consideration,  and  also  the 
air  density. 

Example.  If  the  area  of  a  flat  plate  is  6  square  feet,  the 
co-efficient  K  =  0.003,  and  the  velocity  is  60  miles  per 
hour,  what  is  the  drag  of  the  plate  in  pounds  ?  Solution. 
D  ==  KAV^  =  0.003  X  6  X  (60  X  60)  =  64.80  pounds  drag. 


ELEMENTARY  AERODYNAMICS        77 

For  a  square  flat  plate,  the  co-efficient  K  can  be  taken  as 
0.003. 

Aspect  Ratio.  The  aspect  ratio  of  a  plate  is  the  ratio 
of  the  length  to  the  width.  Thus,  with  an  aspect  ratio  of 
2.0,  we  understand  that  the  plate  is  twice  as  long  as  it  is 
wide.  The  ratio  of  the  length  to  the  width  has  a  very  con- 
siderable influence  of  the  resistance  or  drag,  this  increas- 
ing as  the  ratio  is  made  greater.  If  the  resistance  of  a 
square  plate  is  taken  as  1.00,  the  resistance  of  a  plate  with 
an  aspect  ratio  of  20  will  be  about  1.34  times  as  great. 
The  following  table  will  give  the  effects  of  aspect  ratio 
on  the  resistance  of  a  flat  plane. 

EFFECTS  OF  ASPECT  RATIO  ON  FLAT  PLATES. 

Aspect  Ratio.  Resistance  K  as  a  Multiple 

Length/Width  of  a  Square  Plate. 

1.00  (square)   1.00 

1.50 1.04 

2.00 1.05 

3.00 1.07 

4.00 1.08 

5.00 1.09 

6.00 1.10 

7.00 1.12 

9.00 1.14 

10.00 1.15 

15.00 1.26 

20.00 1.34 

30.00 1.40 

To  convert  the  values  of  a  square  plate  into  a  flat 
plate  of  given  aspect  ratio,  multiply  the  resistance  of  the 
square  plate  by  the  factor  under  the  "K"  heading.  For 
example :  The  resistance  of  a  certain  square  plate  is  20 
pounds,  find  the  resistance  of  a  plate  of  the  same  area, 
but  with  an  aspect  ratio  of  15.    Solution.    The  factor  for 


78        ELEMENTARY  AERODYNAMICS 

a  ratio  of  15  will  be  found  to  be  1.26,  hence  the  resistance 
of  the  required  plate  will  be  20  X  1.26  =  25.2  pounds. 

Streamline  Forms.  When  a  body  is  of  such  form 
that  it  does  not  cause  turbulence  when  moved  through  the 
air,  the  drag  is  entirely  due  to  skin  friction.  Such  a  body 
is  known  as  a  ''streamline  form"  and  approximations  are 
used  for  the  exposed  structural  parts  of  aeroplanes  in 
order  to  reduce  the  resistance.  Streamline  bodies  are  fish- 
like or  torpedo-shaped,  as  shown  by  Fig.  2,  and  it  will  be 


Fig.  2.  Air  Flow  About  a  Streamline  Body  Showing  an  Almost  Complete 
Absence  of  Turbulence  Except  at  the  Extreme  Rear  Edge.  Resist- 
ance Is  Principally  Due  to  Skin  Friction. 


noted  that  the  air  stream  hangs  closely  to  the  outline 
through  nearly  its  entire  length.  The  drag  is  therefore 
entirely  due  to  the  friction  of  the  air  on  the  sides  of  the 
body  since  there  is  no  turbulence  or  "discontinuity."  In 
practical  bodies  it  is  impossible  to  prevent  the  small  tur- 
bulence (I),  but  in  well-designed  forms  its  eflect  is  almost 
negligible. 

In  poor  attempts  at  streamline  form,  the  flow  discon- 
tinues its  adherence  to  the  body  at  a  point  near  the  tail. 
The  poorer  the  streamline,  and  the  higher  the  resistance, 
the  sooner  the  stream  starts  to  break  away  from  the  body 
and    cause    a    turbulent    region.      The    resistance    now 


ELEMENTARY  AERODYNAMICS        79 

becomes  partly  turbulent  and  partly  frictional,  with  the 
resistance  increasing  rapidly  as  the  percentage  of  tlie 
turbulent  region  is  increased. 

The  fact  that  the  resistance  is  due  to  two  factors,  makes 
the  resistance  of  an  approximate  streamline  body  very 
difficult  to  calculate,  as  the  frictional  drag  and  the  tur- 
bulent drag  do  not  increase  at  the  same  rate  for  different 
speeds.  The  drag  due  to  turbulence  varies  as  V-  while 
the  frictional  resistance  only  varies  at  the  rate  of  V^-*"^"', 
hence  the  drag  due  to  turbulence  increases  much  faster 
with  the  velocity  than  the  frictional  component.  If  we 
could  foretell  the  percentage  of  friction,  it  would  be  fairly 
easy  to  calculate  the  total  effect,  but  this  percentage  is 
exactly  what  we  do  not  know.  The  only  sure  method 
is  to  take  the  results  of  a  full  size  test. 

Fig.  2  gives  the  approximate  section  through  a  stream- 
line strut  such  as  used  in  the  interplane  bracing  of  a 
biplane.  The  length  is  (L)  and  the  width  is  (d),  the 
latter  being  measured  at  the  widest  point.  The  relation  of 
the  length  to  the  width  is  known  as  the  "fineness  ratio" 
and  in  interplane  struts  this  may  vary  from  2.5  to  4.5, 
that  is,  the  length  of  the  section  ranges  from  2.5  to  4.5 
times  the  width.  The  ideal  streamline  form  has  a  ratio 
of  from  5.  to  5.75.  Such  large  ratios  are  difficult  to  obtain 
with  economy  on  practical  struts  as  the  small  width  would 
result  in  a  weak  strut  unless  the  weight  were  unduly  in- 
creased. Interplane  struts  reach  a  maximum  fineness 
ratio  at  about  3.5  to  4.5.  Fig.  3  shows  the  result  of  a 
small  fineness  ratio,  the  short,  stubby  body  causing  the 
stream  to  break  away  near  the  front  and  form  a  large  tur- 
bulent region  in  the  rear. 

An  approximate  formula  showing  the  relation  of  fine- 
ness ratio  and  resistance  (curvature  equal)  was  developed 
by  A.  E.  Berriman,  and  published  in  "Flight"  Nov.  12, 
1915.  Let  D  =  resistance  of  a  flat  plate  at  a  given  speed, 
and  R  =  resistance  of  a  strut  at  the  same  speed  and  of 


80 


ELEMENTARY  AERODYNAMICS 


the  same  area,  then  the  relation  between  the  resistance  of 
the  flat  plate,  and  the  strut  will  be  expressed  by  the  formula 
R/D  =  4L/300d,  where  L  =  length  of  section  and  d  = 
width  as  in  Fig.  2.  This  can  be  transposed  for  conven- 
ience, by  assuming  the  drag  of  a  flat  plate  as  D  = 
0.003AV^  where  A  =  area  in  square  feet,  and  V  =  velo- 
city in  miles  per  hour.  The  ratio  of  the  strut  resistance 
to  the  flat  plate  resistance,  given  by  Berriman's  formula, 
can  now  be  multiplied  by  the  flat  plate  resistance,  or  strut 


Fig.  3.  Imperfect  Streamline  Body  with  a  Considerable  Tubulence  Due  to 
the  Short,  Stubby  Form.  Fig.  4  Shows  the  Flow  About  a  Circular 
Rod  or  Cylinder. 

resistance  =  R  =  O.OOSAV^  X  4L/300d.  =  0.012LAVV 
300d.  It  should  be  understood  that  the  area  mentioned 
abov5  is  the  greatest  area  presented  to  the  wind  in  square 
feet,  and  hence  is  equal  to  the  length  of  the  strut  (not 
section)  multiplied  by  the  width  (d). 

Assuming  the  length  (L)  of  the  section  as  7.5  inches, 
and  the  width  (d)  as  one  inch,  the  fineness  ratio  will  be 
7.5.  Using  the  Berriman  formula  in  its  original  form, 
the  relative  resistance  of  the  strut  and  flat  plate  of  same 
area  will  be  found  as  R/D  =  4L/300d  =  0.1,  that  is,  the 
resistance  of  a  streamline  form  strut  of  above  fineness 
ratio  will  be  about  0.1  of  a  flat  plate  of  the  same  area.  It 
should  be  understood  that  this  is  only  an  approximate 


ELEMENTARY  AERODYNAMICS        81 

formula  since  even  struts  of  the  same  fineness  vary  among 
themselves  according  to  the  outline.  Results  published  by 
the  National  Physical  Laboratory  show  streamline  sec- 
tions giving  0.07  of  the  resistance  of  a  flat  plate  of  the 
same  area,  with  fineness  ratio  =  6.5.  In  Fig.  4  the  effects 
of  flow  about  a  circular  rod  is  shown,  a  case  where  the 
fineness  ratio  is  1.  The  stream  follows  the  body  through 
less  than  one-half  of  its  circumference,  and  the  turbulent 
region  is  very  large ;  almost  as  great  as  with  the  flat  plate. 
A  circular  rod  is  far  from  being  even  an  approach  to  a 
perfect  form. 

In  all  the  cases  shown.  Figs.  1-2-3-4,  it  will  be  noticed 
that  the  air  is  affected  for  a  considerable  distance  in  front 
of  the  plane,  as  it  rises  to  pass  over  the  obstruction  before 
it  actually  reaches  it.  The  front  compression  may  be  per- 
ceptible for  6  diameters  of  the  object.  Fi-om  the  examina- 
tion of  several  good  low-resistance  streamline  forms  it 
seems  that  the  best  results  are  obtained  with  the  blunt 
nose  forward  and  the  thin  end  aft.  The  best  position  for 
the  point  of  greatest  thickness  lies  from  0.25  to  0.33  per 
cent  of  the  length  from  the  front  end.  From  the  thickest 
part  it  tapers  out*  gradually  to  nothing  at  the  rear  end. 
That  portion  to  the  rear  of  the  maximum  width  is  the 
most  important  from  the  standpoint  of  resistance,  for  any 
irregularity  in  this  region  causes  the  stream  to  break 
away  into  a  turbulent  space.  From  experiments  it  has 
been  found  that  as  much  as  one-half  of  the  entering  nose 
can  be  cut  away  without  materially  increasing  the  resist- 
ance. The  cut-off  nose  may  be  left  flat,  and  still  the  loss 
is  only  in  the  neighborhood  of  5  per  cent. 

Resistance  Calculations  (Turbulency).  In  any  plate  or 
body  where  the  resistance  is  principally  due  to  turbulent 
action,  as  in  the  flat  plate,  sphere,  cone,  etc.,  the  resistance 
can  be  computed  from  the  formula  R  =  KAV^,  where  R 
is  the  resistance  in  pounds  and  K,  A,  and  V  are  as  before. 
The  resistance  co-efficient  (K)  depends  upon  the  shape  of 


82 


ELEMENTARY  AERODYNAMICS 


the  object  under  standard  air  conditions,  and  differs 
greatly  with  flat  plates,  cones,  sphere,  etc.  The  area  (A) 
is  the  area  presented  to  the  wind,  or  is  the  greatest  area 
that  faces  the  wind,  and  is  taken  at  right  angles  to  its 
direction.  The  following  table  gives  the  value  of  K  for 
the  more  common  forms  of  objects.  See  Figs.  4  to  12, 
inclusive : 


NAME    OF    OBJECT 


(4)  Sphere    

(5)  Hemisphere-Shell  . 

(6)  Hemisphere-Shell  . , 

(7)  Cone-Closed  Base., 

(60°  Angle) 

(8)  Cone-Closed  Base. 

(30°  Angle.) 

(9)  Cone-Hemisphere  . . 

on  Base  (20°) 

(10)  Cone  as  Above 

(11)  Flat  Circular  Disc 

(12)  Flat  Square  Plate, 


F.\CE   OR   EDGE   FACIXG 
WIND 


x\ny  Direction.  . .  . 
Flat  Face  Front.  . 
Round  Face  Front 
Apex  to  Front. . . . 


Apex  to  Front 
Apex  to  Front 


Sphere  to  Front, 
Face  to  Front.  . 
Face  to  Front.  . 


K      IX     TERMS     OP 

Sy.    FT./ MILES 

PER    HR. 


0.000445 
0.003840 
0.008100 
0.001300 

0.000850 

0.000406 

0.0O0222 
0.002820 
0.003000 


The're  are  almost  an  infinite  number  of  dift'erent  forms, 
but  for  the  present  the  above  examples  will  fill  our  pur- 
pose. As  an  example  in  showing  how  greatly  the  form  of 
an  object  influences  its  resistance,  we  will  work  out  the 
resistance  of  a  flat  plate  and  a  spherical  ended  cone,  both 
having  the  same  presented  diameter.  The  cone  is  placed 
so  that  the  spherical  end  will  face  the  air  stream.  The 
area  A  of  both  objects  will  be:  0.7854  X  2  X  2  =  3.1416 
square  feet.  With  an  assumed  wind  velocity  of  100  miles 
per  hour,  the  resistance  of  the  circular  flat  disc  will  be: 
R=  KAV2  =  0.00282  X  3.1416  X  (lOOX  100)  =87.96  lbs. 
For  the  cone,  R  =  KAV2  =  0.000222  X  3.1416  X  (100  X 
100)  =6.97  lbs.  From  this  calculation  it  w^ill  be  seen  that 
it  is  advisable  to  surround  the  object  with  a  spherical  cone 
shaped  body  rather  than  to  present  the  flat  surface  to  the 
wind.    In  the  above  table  the  value  of  K  is  given  for  two 


ELEMENTARY  AERODYNAMICS 


83 


positions  of  the  spherical  based  cone,  the  first  is  with  the 
apex  toward  the  wind,  and  the  second  condition  gives  the 
value  with  the  base  to  the  wind.  With  the  blunt  end  for- 
ward, the  resistance  is  about  one-half  that  when  the 
pointed  apex  enters  the  air  stream.  This  is  due  to  the 
taper  closing-  up  the  stream  without  causing  turbulence. 


K=aooo4^5 


Fig.  6 

K^O.OOQl 


ris.9 

K=O.OOOW6 


rre./a 


Note!  \a/indi5FROm  left  tor/ght-se-e  ar/rows. 


Figs.  4a-5-6-7-8-9-lG-ll-12.  The  Values  of  the  Resistance  Co-efficient  K  for 
Different  Forms  and  Positions  of  Solid  Objects. 
Arrows  Indicate  the  Direction  of  the  Relative 
Wind.      (Eiffel.) 

With  the  apex  forward  there  is  nothing  to  fill  up  the 
vacuous  space  when  the  air  passes  over  the  large  diameter 
of  the  base  as  the  curve  of  the  spherical  end  is  too  short 
to  accomplish  much  in  this  direction. 

Skin  Friction.  The  air  in  rubbing  over  a  surface  expe- 
riences a  frictional  resistance  similar  to  water.  At  the 
present  time  the  accepted  experiments  are  those  of  Dr. 


84        ELEMENTARY  AERODYNAMICS 

Zahm  but  these  are  still  in  some  question  as  to  accuracy. 
It  was  found  in  these  experiments  that  there  was  prac- 
tically no  difference  caused  by  the  material  of  the  surfaces 
as  long  as  they  were  equally  smooth.  Linen  or  cotton 
gave  the  same  results  as  smooth  wood  or  zinc  as  long  as 
there  was  no  nap  or  lint  upon  the  surface.  With  a  fuzzy 
surface  the  friction  increased  rapidly.  This  is  undoubtedly 
due  to  a  minute  turbulence  caused  by  the  uneven  surface, 
and  hence  the  increase  was  not  purely  frictional,  but  also 
due  to  turbulence.  In  the  tests,  the  air  current  was  led 
parallel  to  the  surface  in  such  a  way  that  only  the  friction 
could  move  the  surface.  The  surface  was  freely  sus- 
pended, and  as  the  wind  moved  it  edgewise,  the  movement 
was  measured  by  a  sharp  pointer.  End  shields  prevented 
impact  of  the  air  on  the  end  of  the  test  piece  so  that  there 
was  no  error  from  this  source. 

The  complete  formula  given  by  Dr.  Zahm  is  rather 
complicated  for  ordinary  use,  especially  for  those  not 
used  to  mathematical  computations.  If  Rf  =  resistance 
due  to  friction  on  one  side  of  surface,  L^  length  in  direc- 
tion of  wind  in  feet,  b  =  width  of  surface  in  feet,  and 
V  =  velocity  in  feet  per  second,  then 

Rf  =  0.00000778L«-«^V^-«^b. 

It  will  be  noted  that  the  resistance  increases  at  a  lower 
rate  than  the  velocity  squared,  and  at  a  less  rate  than  the 
area.  That  is  to  say,  that  doubling  the  area  will  not 
double  the  resistance,  but  will  be  less  than  twice  the 
amount.  Giving  the  formula  in  terms  of  area  and  miles 
per  hour  units,  we  have :  Rf  =  0.0000167 A«-«^V^-«^  Where 
A  =  area  in  square  feet  and  V smiles  per  hour.  The 
area  is  for  one  side  of  the  surface  only.  A  rough  approxi- 
mation to  Zahm's  equation  has  been  proposed  by  a  writer 
in  "Flight,"  the  intention  being  to  avoid  the  complicated 
formula  and  yet  come  close  enough  to  the  original  for 
practical   purposes.     The    latter    formula    reads :   Rf  = 


ELEMENTARY  AERODYNAMICS        85 

0.000009V2  where  Rf  and  V  are  as  above.  Up  to  40  miles 
per  hour  the  results  are  very  close  to  Zahm's  formula,  and 
are  fairly  close  from  60  to  90  miles  per  hour.  This  approxi- 
mation is  only  justified  when  the  length  in  the  direction 
of  the  wind  is  nearly  equal  to  the  length.  If  the  length 
is  much  greater,  there  is  a  serious  error  introduced. 

This  formula  is  applied  to  surfaces  parallel  to  the  wind 
such  as  the  sides  of  the  body,  rudder,  stabilizer,  and 
elevator  surfaces  (when  in  neutral).  A  second  important 
feature  of  the  friction  formula  is  that  it  illustrates  the 
law  of  "similitude"  or  the  results  of  a  change  in  scale 
and  velocity,  hence  it  outlines  what  we  must  expect  when 
we  compute  a  full  size  aeroplane  from  the  results  of  a 
model  test. 

The  Inclined  Plane.  When  a  flat  plate  is  inclined  with 
the  wind,  the  resistance  or  drag  will  be  broken  up  into 
two  components,  one  at  right  angles  to  the  air  stream, 
and  one  parallel  to  it.  If  the  plate  is  properly  inclined, 
the  right  angled  component  can  be  utilized  in  obtaining 
lift  as  with  an  aeroplane  wing.  This  is  shown  in  Fig.  13 
where  L  is  the  vertical  lift  force  at  right  angles  to  the  air 
stream  and  D  is  the  horizontal  drag  acting  in  the  direc- 
tion of  the  wind.  As  in  the  case  of  the  plate  placed  normal 
to  the  wind,  there  is  pressure  at  the  front  of  the  plate  and 
a  partial  vacuum  behind.  The  resultant  force  will  be 
determined  by  the  difference  in  pressure  between  the 
front  and  the  back  of  the  plate.  The  forces  will  vary  as 
V^  since  the  reaction  is  caused  by  turbulent  flow.  Both 
the  lift  and  drag  will  vary  with  the  angle  made  with  the 
stream,  and  there  will  be  a  different  value  for  the  co-efli- 
cient  K  for  each  change  in  the  angle.  The  angle  made 
with  the  air  stream  is  known  as  the  "Angle  of  incidence" 
or  the  "Angle  of  attack."  The  change  of  drag  and  lift 
does  not  vary  at  a  regular  rate  with  the  angle. 

A  line  OR  is  the  resultant  of  the  lift  and  drag  forces 
L  and  D,  this  resultant  being  the  force  necessary  to  bal- 


m 


ELEMENTARY  AERODYNAMICS 


ance  the  two  forces  L-D.  It  is  on  the  point  of  applica- 
tion O  that  the  plate  balances,  and  this  point  is  sometimes 
known  as  the  "Center  of  pressure."  The  center  of  pres- 
sure is  therefore  the  point  at  which  the  resultant  inter- 
sects the  surface  of  the  flat  plate.  The  resultant  OR  is 
approximately  at  right  angles  to  the  surface  at  small 
incident  angles,  and  the  point  O  is  nearer  the  front  or 
^'leading  edge"  (A)  of  the  plane.  The  smaller  the  angle 
of  incidence  the  nearer  will  the  point  O  approach  the  lead- 
ing edge  A.     By  drawing  OL  to  scale,  representing  the 


NOfSMAL  PLAN£. 
(C.P  C£/^77SACi 


ir^CUNEDPLAtli 


CC.RN£Af2L^ 


Fig.  13.  Flow  About  Inclined  Plane  and  Forces  Produced  by  Stream.  Fig. 
14.  Normal  Plane  with  C.P.  at  center  of  Plate.  Fig.  15.  C.P. 
Moves  Toward  Entering  Edge  When  Plate  Is  Inclined  to  Wind. 

lift,  and  OD  to  scale  representing  the  drag,  we  can  find 
the  resultant  OR  by  drawing  LR  parallel  to  the  drag  OD 
and  DR  parallel  to  the  lift  line  OL.  All  lines  drawn 
through  the  intersection  of  LR  and  DR  will  give  the 
resultant  OR  to  scale.  All  of  the  lines  must  be  started 
from  the  center  of  pressure  at  O. 

The  least  resultant  will,  of  course,  occur  Avhen  the  plane 
is  parallel  to  the  air  stream.  The  maximum  resultant 
will  occur  when  the  angle  of  incidence  is  about  40  degrees, 


ELEMENTARY  AERODYNAMICS        87 

and  on  a  further  increase  in  the  angle,  the  value  of  the 
resultant  will  gradually  decrease.  When  the  plane  is 
parallel  with  the  stream,  the  resultant  is  parallel  to  the 
plate,  but  rapidly  approaches  a  position  at  right  angles 
at  about  an  incidence  of  6  to  10  degrees.  Beyond  10 
degrees  incidence  the  angle  of  the  resultant  increases  past 
the  normal. 

The  center  of  pressure  (O),  or  the  point  where  the  re- 
sultant force  intersects  the  plane,  moves  forward  as  the 
angle  of  incidence  is  decreased  from  90°.  When  at  right 
angles  to  the  air  current,  the  center  of  pressure  is  exactly 
in  the  center  of  the  plane  as  shown  by  Fig.  14.  In  this  case 
the  drag  (D)  is  the  resultant,  and  acting  in  the  center,  ex- 
actly balances  the  air  forces.  In  Fig.  15  the  angle  of  inci- 
dence is  reduced,  consequently  the  center  of  pressure 
moves  nearer  the  leading  edge  (A).  As  the  angle  con- 
tinues to  decrease,  the  C.  P.  moves  still  further  forward 
until  it  lies  directly  on  the  front  edge  when  the  plate  be- 
comes parallel  with  the  air  stream.  The  center  of  pressure 
movement  is  due  to  the  fact  that  more  and  more  work  is 
done  by  the  front  part  of  the  surface  as  the  angle  is 
decreased.  Consequently  the  point  of  support,  or  C.  P., 
must  move  forward  to  come  under  the  load.  It  should 
be  understood  that  the  plane  will  balance  about  the  C.  P. 
if  a  knife  edge  bearing  were  applied  as  at  R  in  Fig.  15. 

Calculation  of  Inclined  Planes.  We  will  now  consider 
the  inclined  plane  as  a  lifting  surface  for  an  aeroplane, 
and  make  the  elementary  calculations  for  such  purpose. 
The  lift  will  first  be  calculated  for  the  support  of  the 
given  load,  at  the  given  velocity,  and  then  the  drag.  For 
several  reasons,  that  will  afterwards  be  explained,  the 
flat  plate  or  plane  is  not  used  for  the  main  lifting  surfaces, 
but  the  experience  gained  in  computing  the  plate  will  be 
of  great  assistance  when  we  start  calculating  actual  wings. 

Lift  and  Drag  Co-efihcients.  The  lift  component  (L)  of 
the  inclined  flat  plate  depends  on  the  velocity,  area,  aspect 


88       ELEMENTARY  AERODYNAMICS 

ratio  and  angle  of  incidence.  Instead  of  using  the  co-effi- 
cient (K)  formerly  used  for  the  total  drag,  we  will  use  the 
lift  co-efficient  Ky.  The  formula  for  lift  now  becomes: 
L=KyAV^  where  A  =  area  in  square  feet,  and  V  = 
velocity  in  miles  per  hour.  The  lift  co-efficient  Ky  depends 
upon  the  angle  of  incidence.  The  horizontal  drag  D  will 
be  calculated  from  the  drag  co-efficient  Kx  which  is  used 
in  the  same  way  as  the  co-efficient  K  in  the  case  of  the 
normal  plate.  The  subscript  (x)  is  used  to  distinguish  it 
from  the  lift  co-efficient.  Both  Ky  and  Kx  must  be  cor- 
rected for  aspect  ratio.  The  drag  can  be  calculated  from 
the  formula :  D  =  KxAV^  where  the  letters  A  and  V  are 
the  same  as  above. 

For  the  calculation  of  the  drag,  we  will  use  a  new 
expre,ssion — the  "Lift-Drag  Ratio" — or  as  more  com- 
monly given,  "L/D."  This  shows  the  relation  between 
the  lift  and  drag,  so  that  by  knowing  the  lift  and  the  ratio 
for  any  particular  case,  we  can  compute  the  drag  without 
the  necessity  of  going  through  the  tedious  calculation 
D  =  KxAV^.  The  lift-drag  ratio  for  a  flat  plate  varies 
with  the  angle  of  incidence,  and  the  aspect  ratio,  and  hence 
a  separate  value  must  be  used  for  every  inclination  and 
change  in  aspect.  To  obtain  the  drag,  divide  the  lift  by 
the  lift-drag  ratio.  Hence  if  the  lift  is  1200  pounds,  and 
the  ratio  equals  6.00,  the  drag  will  be:  1200/6  =  200 
pounds,  or  in  other  words,  the  lift  is  6  times  the  drag 
force.  Changing  the  angle  of  incidence  through  angles 
ranging  from  1  degree  to  7  degrees,  the  lift-drag  ratio  of 
a  flat  plate  will  vary  from  1.5  to  7.5.  When  the  plane  is 
parallel  to  the  wind  stream  and  gives  no  lift,  the  drag 
is  computed  from  Zahm's  skin  friction  formula. 

The  following  tables  give  the  values  of  Ky,  Kx,  L/D, 
and  center  of  pressure  movement  for  flat  plates  of  various 
aspect  ratios.  The  center  of  pressure  (C.  P.)  for  each 
angle  is  given  as  a  decimal  fraction  of  its  distance  from 
the  leading  edge,  in  terms  of  the  width  or  "Chord." 


ELEMENTARY  AERODYNAMICS 


89 


Fig.  16  shows  the  top  view  or  plan  of  a  lifting  surface, 
with  the  direction  of  the  wind  stream  indicated  by  the 
arrows  w-w-w  =  w.  The  longer  side  or  "span"  is  indicated 
by  S,  while  the  width  or  chord  is  C.  Main  lifting  surfaces, 
or  wings,  have  the  long  side  at  right  angles  to  the  wind 
as  shown.    When  in  this  position,  the  surface  is  said 


Fig.  16.  Plan  View  of  Plate  with  Long  Edge  to  Wind.  Fig.  17.  Plate 
with  Narrow  Edge  to  Wind,  Showing  Loss  in  Lift.  17a  Shows 
Effect  of  Raked  Tips. 

to  be  in  "Pterygoid  Aspect,"  and  when  the  narrow  edge 
is  presented  to  the  wind,  the  wing  is  in  "Apteroid  Aspect." 
The  word  "Pterygoid"  means  "Bird  like,"  and  was  chosen 
for  the  condition  in  Fig.  16,  as  this  is  the  method  in  which 
a  bird's  wing  meets  the  air.  Contrary  to  the  case  with 
true  curved  aeroplane  wings,  flat  planes  usually  give 
better  lift  in  apteroid  than  in  pterygoid  aspect  at  high 
angles.  The  aspect  ratio  will  be  the  span  (S)  divided  by 
the  chord  (C),  or  Aspect  ratio  =  S/C. 


90 


ELEMENTARY  AERODYNAMICS 


AERODYNAMIC   PROPERTIES   OF    INCLINED    FLAT   PLATES. 


ASPECT  RATia=l. 


ASPECT  RATIO  =  1.3. 


ANGLE 
OF  IN- 
CIDENCE 
5 

10 
20 
30 
45 


1 

1 

1     L 

1 

L 

i    LIFT 

DRAG 

— 

LIFT 

DRAG 

Ky 

Kx 

D 

C.P. 

Ky 

Kx 

D 

0.00045 

0.00007 

6.3 

0.231 

0.00057 

0.00011 

6.5 

.00097 

.00019 

5.1 

0.268 

.00109 

.00020 

5.6 

.00208 

.00074 

2.8 

.328 

.00215 

.00077 

2.8 

.00294 

.00173 

1.7 

.392 

.00198 

.00095 

1.7 

.00207 

.00210 

1.0 

.432 

.00185 

.00184 

1.2 

C.P. 

0.230 
.271 
.331 
.398 

.442 


ASPECT  RATIO  = 

2. 

1 

ASPECT  RATIO  = 

3. 

5 

0.00069 

0.00009 

7.8 

0.231 

0.00087 

0.00012 

7.7 

0.233 

10 

.00123 

.00021 

5.9 

0.278 

.00140 

.00028 

5.1 

.300 

20 

.00247 

.00091 

2.7 

0.345 

.00210 

.00077 

2.7 

.406 

30 

.00178 

.00111 

1.7 

0.405 

.00193 

.00111 

1.7 

.420 

40 

.00169 

.00146 

1.2 

0.439 

.00169 

.00122 

1.1 

.426 

ASPECT  RATIO  = 

6. 

II             ASPECT  RATIO  = 

9. 

5 

0.00103 

.00016 

6.4 

0.275 

0.00134           .00026         5.2 

0.289 

10 

.00173 

.00034 

5.2 

.333 

.00186           .00040         4.7 

0.348 

20 

.00199 

.00074 

2.7 

.391    1 

.00211     1      .00080    1    2.6    1 

.0.390 

30 

.00204 

.00120 

1.7 

.406 

1      .00210    1      .00127    1     1.7 

0.398 

40 

.00191 

.00116 

1.1 

.421     1      .00194    1      .00179    1     1.1 

0.416 

It  will  be  seen  from  the  above  that  the  lift  coefficient 
Ky  increases  with  the  aspect  ratio,  and  that  it  generally 
declines  after  an  angle  of  30  degrees.  The  center  of  pres- 
sure moves  steadily  back  with  an  increase  in  angle. 

Example  for  Lifts.  A  certain  flat  plane  has  an 
area  of  200  square  feet,  and  moves  at  50  miles  per  hour. 
The  angle  of  incidence  is  10  degrees,  and  the  aspect 
ratio  is  6.  Find  the  total  lift  and  the  drag  in  pounds.  Also 
the  location  of  the  center  of  pressure  in  regard  to  the 
leading  edge,  if  the  chord  is  5.8  feet. 

Solution.  Under  the  table  headed,  "Aspect  Ratio 
=  6  we  find  that  Ky  at  10°  ==0.00173,  and  that  the  lift 
drag  ratio  is  5.2.  The  center  of  pressure  is  0.333  of  the 
chord  from  the  front  edge.  The  total  lift  then  becomes : 
L=KyAV2  =  0.00173x200x( 50x50)  =865  pounds.  Since 
the  lift  drag  ratio  is  5.2,  the  drag=  D  =  865/5.2=  166.3 
pounds.  The  center  of  pressure  will  be  located  5,8x0.333 
=  1.4  feet  from  the  leading  edge. 

Under  the  same  conditions,  but  with  an  aspect  ratio 
of  3,  the  Hft  will  become :  L  =  KyAV2  =  0.0014x200x(50x 
50)  =  700  pounds.  In  this  case  the  lift  drag  ratio  is  5.1, 
so  that  the  drag  will  be  137.8  pounds.  Even  with  the 
same  area,  the  aspect  ratio  makes  a  difference  of  865 — 700 


ELEAIENTARY  AERODYNAMICS  91 

=  165  pounds.  If  we  were  compelled  to  carry  the  original 
865  pounds  with  aspect  3  wing,  we  would  also  be  com- 
pelled to  increase  the  area,  angle,  or  speed.  If  the  speed 
were  to  be  kept  constant,  we  would  be  limited  to  a  change 
in  area  or  angle.  In  the  latter  case  it  would  be  preferable 
to  increase  the  area,  since  a  sufficient  increase  in  the  angle 
would  greatly  increase  the  drag.  It  will  be  noted  that  the 
lift-drag  ratio  decreases  rapidly  with  an  increase  in  the 
angle. 


Burgess  Seaplane  Scout, 

Calculation  of  Area:  Let  us  assume  that  we  are 
confined  to  the  use  of  an  aspect  ratio  of  6,  a  speed 
of  50  miles  per  hour,  weight  =  2500  pounds,  and  wish 
to  obtain  the  area  that  will  give  the  most  efficient  surface 
(Least  lift-drag  ratio.)  The  equation  can  be  now  trans- 
posed so  that  the  area  =  A==KyV^  On  examination 
of  the  table  it  will  be  seen  that  the  greatest  lift-drag 
ratio  is  6.4  at  5  degrees,  and  that  the  Ky  at  this  angle  is 
0.00103.  Substituting  these  values  in  the  equation  for 
area,  we  have  A  =  L/Ky V^  =  2500/0.00103  x  (50  x  50)  = 
971  square  feet. 


92 


LABORATORIES 


Wind  Tunnel  at  Washington  Navy  Yard  in  Which  the  Air  Circulates  Con- 
tinuously Through  a  Closed  Circuit. 


CHAPTER  IV. 
EXPERIMENTAL  LABORATORIES. 

Test  Methods  in  General.  As  already  explained,  the 
behavior  of  a  body  in  an  air  stream  cannot  be  predicted 
with  any  certainty  by  direct  mathematical  calculation, 
and  for  this  reason,  each  and  every  aerodynamic  body 
must  be  tested  under  conditions  that  are  as  nearly  similar 
to  the  actual  working  conditions  as  possible.  Prior  to 
Professor  Langley's  first  experiments  in  1887,  mechan- 
ical flight  with  a  heavier  than  air  machine  was  derided  as 
an  impossibility,  even  by  such  scientists  as  Navier,  Von 
Helmotz,  Gay-Lussac,  and  others,  who  proved  by  the 
most  intricate  calculations  that  a  body  larger  than  a  bird 
could  not  be  supported  by  its  own  energy.  Such  calcula- 
tions were,  of  course,  based  on  a  wrong  understanding  of 
air  flow,  and  as  no  experimental  w^ork  had  been  done  up 
to  that  time,  the  flow  was  assumed  according  to  the  indi- 
vidual taste  and  belief  of  the  demonstrator.  The  presence 
of  a  vacuum  on  the  back  of  a  plate  was  not  understood, 
and  as  this  contributes  full  two-thirds  of  the  lift,  it  is  an 
easy  matter  to  see  why  all  of  the  early  predictions  fell 
short  of  the  actual  lifting  forces.  To  quote  one  classic 
absurdity,  the  scientist  Navier  proved  mathematically 
that  if  mechanical  flight  were  possible,  then  17  swallows 
would  be  capable  of  developing  one  horsepower. 

In  spite  of  these  discouraging  computations,  Langley 
proceeded  with  a  very  carefully  conducted  series  of  expe- 
riments, first  investigating  the  laws  of  surface  sustena- 
tion  on  various  forms  of  plates,  and  when  the  data 
collected  was  sufficient  for  his  needs,  he  started  to  con- 

93 


94  LABORATORIES 

struct  a  number  of  model  flyers  with  various  wing 
arrangements  and  aerofoil  forms.  It  was  Langley's  expe- 
riments upon  aerofoils  that  cleared  the  way  for  the 
Wright  Brothers,  who  started  a  further  and  more  com- 
plete investigation  in  1896.  Experiments  were  made  on 
the  effect  of  curvature,  aspect  ratio  and  angle  of  inci- 
dence, and  the  results  obtained  in  their  "wind  tunnel" 
were  afterwards  applied  to  their  successful  full  size  ma- 
chine. During  1901  to  1902  the  Wrights  investigated  the 
properties  of  at  least  100  different  aerofoil  forms.  Both 
Langley's  and  Wrights'  experiments  were  with  models, 
although  they  were  made  in  a  different  manner.  It  was 
in  this  way  that  experimental  evidence  gained  precedence 
over  theory. 

Langley's  specimens  were  mounted  at  the  end  of  a 
revolving  arm,  so  that  with  the  arm  revolving,  a  relative 
air  stream  of  known  velocity  could  be  had.  The  aero- 
foil was  mounted  in  such  a  way  that  the  lift  and  drag 
could  be  measured.  In  the  early  experiments  of  the 
Wrights,  the  models  were  placed  in  an  enclosed  channel 
through  which  a  stream  of  air  was  maintained  by  a  fan. 
The  model  was  attached  to  a  balance  system  so  that  the 
lift  and  resistance  could  be  measured.  This  is  what  is 
now  known  as  a  "wind  tunnel,"  and  at  present  is  almost 
exclusively  used  in  model  tests.  Several  investigators 
immersed  their  model  aerofoils  in  running  water  so  that 
the  direction  of  flow  could  be  visibly  observed.  While 
this  latter  method  is  of  great  service  in  determining  dis- 
turbances, stream  line  flow,  and  general  characteristics, 
it  is  qualitative  rather  than  quantitative,  and  cannot  be 
used  in  obtaining  accurate  numerical  results.  A  more 
accurate  method  of  mapping  out  the  direction  of  flow, 
eddies,  etc.,  is  to  introduce  smoke  into  the  air  stream. 

Full  Size  Experiments.  The  old  "rule  of  the  thumb" 
method  of  building  a  full  size  machine  without  model  test 
data  or  other  experimental  evidence  to  begin  with  has 


LABORATORIES  95 

seen  its  day.  It  is  not  only  exceedingly  expensive,  but 
is  highly  dangerous,  and  many  a  flyer  has  met  his  death 
in  the  endeavor  to  work  out,  untried  principles  on  a  full 
size  machine.  The  first  cost  of  the  machine,  the  continual 
breakage  and  operating  expense,  to  say  nothing  of  the 
damage  suits  and  loss  of  time,  make  a  preliminary  full  size 
tryout  an  absurdity  at  the  present  time.  Again,  the  results 
of  full  size  experiments  are  not  always  reliable,  as  so 
much  depends  upon  the  pilot  and  weather  conditions. 
The  instruments  used  on  a  large  machine  are  far  from 
being  as  accurate  as  those  used  in  model  tests.  These 
are  also  likely  to  be  thrown  out  of  adjustment  unknow- 
ingly by  falls  or  collisions.  The  great  number  of  variables 
that  enter  into  such  a  test  make  it  almost  an  impossibility 
to  obtain  accurate  data  on  the  result  of  minor  alterations, 
and,  in  fact,  it  is  almost  impossible  to  get  the  same  results 
twice  without  further  alterations  than  changing  the  pilot. 
Full  scale  tests  are  necessary  after  sufficient  data  has 
been  obtained  and  applied  in  a  scientific  manner  to  the 
design  of  the  machine,  but  successful  performance  can- 
not be  expected  from  a  powered  machine  built  by  guess 
work. 

When  performed  in  connection  with  a  wind  tunnel,  or 
based  on  dependable  data  from  other  sources,  full  size 
wing  tests  are  very  instructive  and  useful  if  care  is  taken 
to  have  the  tests  conducted  under  uniform  and  known 
conditions.  Many  full  size  experiments  of  this  nature 
have  been  carried  out  by  Saint-Cyr  University  in  France, 
and  by  the  Royal  Aircraft  Factory  in  Great  Britain.  Both 
of  these  institutions  have  a  wind  tunnel  and  an  almost 
unlimited  fund  of  performance  data,  and  last  but  not 
least,  have  the  services  of  skilled  observers. 

At  Saint-Cyr,  the  full  size  wings,  or  the  entire  machine, 
are  carried  on  an  electric  car  or  "chariot."  The  speed  of 
the  car,  the  lift  and  drag,  can  be  determined  at  any  mo- 
ment during  the  run  through  suitable  recording  devices. 


96  LABORATORIES 

Actual  flying  tests  have  also  been  made,  the  measure- 
ment of  the  propeller  thrust  giving  the  drag,  while  the 
lift  is  known  as  being  equal^to  the  weight  of  the  machine. 
The  R.  A.  F.  have  carried  out  a  very  extensive  series  of 
flight  tests,  the  experiments  on  the  old  "B.  E.-2"  probably 
being  the  best  known. 

The  greater  part  of  the  experiments  performed  with  the 
car  at  Saint-Cyr  differed  considerably  from  the  results 
obtained  by  model  tests,  and  apparently  these  differences 
followed  no  specific  law.  According  to  theory,  and  the 
results  obtained  by  different  laboratories,  the  performance 
of  a  full  size  wing  should  be  better  than  with  a  model, 
but  the  Saint-Cyr  tests  showed  that  such  was  not  always 
the  case.  The  center  of  pressure  movement  differed  in 
almost  every  case,  and  as  a  direct  result,  the  pressure 
distribution  of  the  large  wings  was  materially  different 
than  with  the  model.  The  lift-drag  ratio  results  varied, 
sometimes  being  better  for  the  model  than  for  the  large 
wing.  These  differences  can  probably  be  explained  as 
being  due  to  variation  in  air  currents,  side  winds,  etc. 

Model  Tests.  Since  lift  and  resistance  are  due  to  rela- 
tive motion  between  a  body  and  the  air  stream,  a  model 
can  either  be  towed  through  the  air,  or  it  can  be  held 
stationary  while  the  air  is  forced  past  it.  There  has  been 
some  controversy  on  the  relation  between  the  results 
obtained  by  the  two  methods,  but  for  the  present  we  will 
accept  the  common  belief  that  the  results  obtained  by 
either  method  are  the  same.  In  testing  ship  models, 
they  are  always  towed  through  the  tank,  but  in  the  case  of 
aero-dynamic  bodies  this  is  complicated  and  not  desir- 
able. In  towing  models  through  the  air  a  very  high 
velocity  is  needed  and  this  necessitates  either  a  very  long 
track  or  a  short  time  length  for  making  the  observations. 
Again,  it  is  almost  impossible  to  avoid  errors  because 
of  vibration,  inequality  of  movement  due  to  uneven  track, 
or  air  eddies  caused  by  differences  in  temperature  and  by 


LABORATORIES  97 

the  movement  of  the  towing  device.  In  fact,  the  same 
difficulties  apply  to  towed  model  tests  as  to  the  full  size 
"electric  chariot." 

The  whirling  arm  method  of  testing  as  used  by  Lang- 
ley,  Maxim-Vickers,  and  others,  is  a  form  of  ''towed 
testing,"  but  is  also  open  to  serious  objections.  Unless 
the  arm  is  very  long,  every  part  of  the  model  surface  will 
not  move  at  the  same  velocity,  the  outer  portions  moving 
the  faster.  As  the  forces  produced  by  an  air  stream  vary 
as  the  square  of  the  velocity,  this  may  introduce  a  serious 
error.  The  fact  that  the  body  passes  repeatedly  over  the 
same  path  introduces  error,  as  the  body  after  the  first 
revolution  is  always  working  in  disturbed  air.  The  centri- 
fugal force,  and  the  currents  set  up  by  the  arm  itself  all 
reduce  the  accuracy  of  the  method. 

When  a  model  is  placed  in  a  uniform  current  of  air  in 
a  properly  designed  channel  or  tunnel,  the  greater  part 
of  the  errors  due  to  towed  tests  are  eliminated.  The 
measuring  instruments  can  be  placed  on  a  firm  founda- 
tion, the  air  stream  can  be  maintained  at  a  nearly  uniform 
speed  and  with  little  error  due  to  eddies,  and  the  test  may 
be  continued  under  uniform  conditions  for  an  indefinite 
period.  While  there  are  minor  errors  due  to  wall  fric- 
tion and  slight  variations  in  the  velocity  at  different 
points  in  the  cross  section  of  the  tunnel,  they  are  very 
small  when  compared  to  the  errors  of  towing.  For  this 
reason  the  wind  tunnel  is  the  accepted  means  of  testing. 

Eiffel's  Wind  Tunnel.  The  Eiffel  Laboratory  at 
Auteuil,  France,  is  probably  one  of  the  best  known.  The 
results  in  Chapters  III  and  V  were  obtained  in  this 
laboratory  and  thousands  of  similar  experiments  have 
been  carried  out  at  this  place.  Two  tunnels,  a  large  and 
small,  are  placed  side  by  side  in  the  main  laboratory  room, 
the  tunnels  being  supported  midway  between  the  floor 
and  ceiling.  The  air  is  drawn  from  this  room  into  an  air- 
tight experimental  chamber  through  a  bell-mouthed  cir- 


98  LABORATORIES 

cular  opening.  A  grill  or  honeycomb  baffle  is  placed  in 
the  opening  to  straighten  out  the  flow,  and  from  this 
point  the  air  passes  across  the  chamber  and  exits  through 
a  circular  duct  to  the  suction  side  of  a  large  fan.  From 
the  fan  the  air  is  discharged  into  the  room.  The  same 
air  thus  circulates  through  the  tunnel  continuously.  The 
test  chamber  is  considerably  wider  than  the  openings  so 
that  the  walls  do  not  influence  the  flow  around  the  model. 
A  cylinder  of  air  passes  through  the  chamber  at  a  remark- 
ably uniform  velocity,  and  without  any  appreciable  eddies. 
Diameter  of  the  stream  approximates  6.6  feet  in  the  large 
tunnel  and  3  feet  in  the  smaller.  In  the  large  tunnel  the 
maximum  velocity  is  105  feet  per  second,  and  131  feet 
per  second  is  attained  in  the  smaller.  A  50-horsepower 
electric  motor  is  used  with  a  multiblade  fan  of  the 
"Sirocco"  type. 

The  observer  and  weighing  mechanism  are  supported 
above  the  air  stream  on  a  sliding  floor,  and  a  standard 
extends  from  the  model  in  the  wind  stream  to  the  bal- 
ances on  the  weighing  floor.  These  balances  determine 
the  lift  and  drag  of  the  models,  the  center  of  pressure,  etc. 

The  N.  P.  L.  Tunnel.  The  National  Physical  Labora- 
tory at  Teddington,  England,  has  a  remarkably  com- 
plete and  accurate  aerodynamic  equipment.  This  con- 
sists of  a  large  tunnel  of  7  square  feet  area,  a  small 
tunnel  of  4  square  feet,  and  a  whirling  table  house.  The 
large  tunnel  is  80  feet  in  length  with  an  air  flow  of  60 
feet  per  second,  the  air  being  circulated  by  a  four-bladed 
propeller  driven  by  an  electric  motor  of  30  horsepower. 
The  velocity  is  uniform  within  one-half  per  cent,  and  the 
most  accurate  of  results  have  been  obtained.  The  smaller 
tunnel  is  about  56  feet  long  and  the  wind  velocity  is  about 
40  miles  per  hour  maximum.  The  propeller  revolves  at 
600  revolutions  and  is  driven  by  a  10-horsepower  elec- 
tric motor.  There  is  no  chamber  and  the  models  are 
suspended  in  the  passage  half  way  between  the  "DifTuser" 


LABORATORIES  99 

in  the  entering  end,  and  the  baffles  in  the  exit.  The 
Massachusetts  Institute  of  Technology,  and  the  Curtis 
Aeroplane  Company  both  have  similar  tunnels. 

United  States  Navy  Tunnel.  In  this  tunnel  the  air  is 
confined  in  a  closed  circuit,  the  return  tunnel  being  much 
larger  than  the  section  in  which  the  tests  are  performed. 
The  cross-sectional  area  is  8  square  feet  at  the  point  of 
test,  and  the  stream  is  uniform  within  2  per  cent.  The 
balance  and  controls  are  mounted  on  the  roof  of  the 
tunnel,  with  an  arm  extending  down  through  the  air 
stream  to  the  model,  as  in  the  Eiffel  tunnel.  The  balance 
is  similar  to  Eiffel's  and  is  sensitive  to  less  than  2/1000 
pound.  A  velocity  of  7h  miles  per  hour  may  be  attained 
by  the  500-horsepower  motor,  but  on  account  of  the  heat- 
ing of  the  air  stream  through  skin  friction,  the  tests  are 
generally  made  at  40  miles  per  hour.  Models  up  to  36-inch 
span  can  be  tested,  while  the  majority  of  models  tested 
at  M.  I.  T.  are  about  18  inches. 


ZolW 


CHAPTER  V. 

AERODYNAMICS  OF  LIFTING  SURFACES 
(AEROFOILS). 

General  Wing  Requirements.  The  performance  of  flat 
plates  when  used  as  Hfting  surfaces  is  very  poor  compared 
with  curved  sections  or  wing  forms.  It  will  be  remembered 
that  the  greatest  lift-drag  ratio  for  the  flat  plate  was  6.4, 
and  the  best  Ky  was  0.00294.  Modern  wing  sections  have 
a  lift-drag  ratio  of  over  20.0,  and  some  sections  have  a 
lift  coefficient  of  Ky=  0.00364,  or  about  60  per  cent 
higher  than  the  lift  obtained  with  a  flat  plate.  In  fact, 
this  advantage  made  flight  possible.  To  Langley,  above 
all  other  men,  we  owe  a  debt  of  gratitude  for  his  investi- 
gations into  the  value  of  curved  wing  surfaces. 

Air  Flow  About  an  Aerofoil.  To  distinguish  the  curved 
wing  from  the  flat  plane,  we  will  use  the  term  ''Aerofoil." 
Such  wings  are  variously  referred  to  as  "Cambered  sur- 
faces," "Arched  surfaces,"  etc.,  but  the  term  "Aerofoil"  is 
more  applicable  to  curved  sections.  The  variety  of  forms 
and  curvatures  is  almost  without  limit,  some  aerofoils  be- 
ing curved  top  and  bottom,  while  others  are  curved  only 
on  the  upper  surface.  The  curve  on  the  bottom  face  may 
either  be  concave  or  convex,  an  aerofoil  of  the  latter  type 
being  generally  known  as  "Double  cambered."  The  curves 
may  be  circular  arcs,  as  in  the  Wright  and  Nieuport  wings, 
or  an  approximation  to  a  parabolic  curve  as  with  many 

of  the  modern  wings. 

Fig.  1-b  shows  the  general  trend  of  flow  about  an  aero- 
foil at  two  different  angles  of  incidence,  the  flow  in  the 
upper  view  being  characteristic  for  angles  up  to  about  6 

100 


AEROFOILS 


101 


degrees,  while  the  lower  view  represents  the  flow  at  angles 
approximating  16°.  At  greater  angles  the  air  stream 
breaks  away  entirely  from  the  top  surface  and  produces  a 
turbulence  that  greatly  resembles  the  disturbance  pro- 
duced by  a  flat  plate.  It  will  be  noted  in  the  top  figure 
(At  small  angles)  that  the  flow  is  very  similar  to  the  flow 
about  a  streamline  body,  and  that  the  air  adheres  very 
closely  to  the  top  surface.  The  flow  at  small  angles  is 
very  steady  and  a  minimum  of  turbulence  is  produced  at 
the  trailing  edge. 

When  increased  beyond  6^,  turbulence  begins,  as  shown 
in  the  lower  figure,  and  a  considerable  change  takes  place 


Aerofoil  Types 


Flow  /About  Aerofoils- 


LOWER  FACC  CAM9EH           LOWER  FACE  PLAT 
1  > ^.WIND 


FLOW  AT  /^PPROX.  & 


WRIGHT  TYPE  CURVED  PLATE 

Fig.  No.l-^ 


FLOW  ATAPPROX-  tG° 

FiG.No.l-l> 


Figs,   la,  lb.     Aerofoil    Types    and    Flow    at    Different    Angles. 


in  the  lift-drag  ratio.  This  is  known  as  the  "Lower  Crit- 
ical Angle."  The  turbulence,  however,  is  confined  to  the 
after  part  of  the  wing,  and  little  or  no  disturbance  takes 
place  in  the  locality  of  the  lower  surface.  We  observe 
that  an  increase  in  angle  and  lift  produces  an  increased 
turbulent  flow  about  the  upper  surface,  and  hence  the 
upper  surface  is  largely  responsible  for  the  lift.  Below 
10"  the  trend  of  the  upper  portion  of  the  stream  is  still 
approximately  parallel  to  the  upper  surface. 


102 


AEROFOILS 


From  16°  to  18°,  the  stream  suddenly  breaks  entirely 
away  from  the  wing  surface,  and  produces  an  exceedingly 
turbulent  flow  and  mass  of  eddies.  The  lift  falls  off  sud- 
denly Avith  the  start  of  the  discontinuous  flow.  The  angle 
at  which  this  drop  in  lift  takes  place  is  known  variously 
as  the  ''Second  Critical  Angle,"  the  "Burble  Point,"  or 
the  ''Stalling  Angle."  Any  further  increase  in  angle  over 
the  stalling  angle  causes  a  drop  in  lift  as  the  discontinuity 
is  increased.  With  the  flat  plane,  the  burble  point  occurs  in 
the  neighborhood  of  30°  and  movement  beyond  this  angle 


3r-^ 

y- WIND 

^^^^  ^."-'^ 

_-_-^:_:=r-^ 

.<^^^^^^^ 

^^p^r^g== 

--;;;;^^^^'''C1>?/S//N«S   TREND 

■              LINEOrF-UGHT--^       "^ 

Fig.  2.     Showing  How  Lift  Is  Obtained  When  an  Aerofoil  Is  Inclined  at  a 
Negative  Angle,  the  Line  of  Flight  Being  Along  X-X. 

also  decreases  the  lift.  In  flight,  the  burble  point  should 
not  be  approached,  for  a  slight  increase  in  the  angle  when 
near  this  point  is  likely  to  cause  the  machine  to  drop  or 
"Stall."  The  fact  that  the  maximum  lift  occurs  at  the 
critical  angle  makes  the  drop  in  lift  at  a  slightly  greater 
angle,  doubly  dangerous. 

A  peculiar  feature  of  the  aerofoil  lies  in  the  fact  that 
lift  is  still  obtained  with  a  zero  angle  of  incidence,  and 
even  with  a  negative  angle.  With  the  aerofoil  shown  in 
Fig.  2  there  will  be  a  considerable  lift  when  the  flat  bottom 
is  parallel  with  the  direction  of  travel,  and  some  lift 
will  still  be  obtained  with  the  front  edge  dipped  down 
(Negative  Angle).   The  curved  upper  surface  causes  the 


AEROFOILS  103 

air  stream  to  rise  toward  the  front  edge,  as  at  E,  hence 
the  wing  can  be  dipped  down  considerably  in  regard  to 
the  line  of  motion  X-X,  without  going  below  the  actual 
air  stream. 

Action,  in  Producing  Lift.  At  comparatively  high  angles 
of  incidence,  where  there  is  turbulent  flow,  the  lift  and 
drag  are  due  principally  to  the  difference  in  pressure 
between  the  upper  and  lower  surfaces  as  in  the  case  of 
the  flat  plate. 

There  is  a  positive  pressure  below  as  in  the  front  of  a 
flat  inclined  plane,  and  a  vacuous  region  above  the  upper 
surface.  The  drag  with  the  plane  below  the  burble  point, 
and  above  the  ** Lower  Critical  Angle,"  is  due  both  to  skin 
friction  and  turbulence — principally  to  the  latter.  Below 
the  first  critical  angle  (6°),  the  skin  friction  effect  in- 
creases, owing  to  the  closeness  w'ith  w^hich  the  air  stream 
hangs  to  the  upper  surface. 

Since  there  is  but  little  turbulence  at  the  small  angles 
below  6",  the  theory  of  the  lift  at  this  point  is  difficult  to 
explain.  The  best  explanation  of  lift  at  small  angles  is 
given  by  Kutta's  Vortex  Hypothesis.  This  theory  is  based 
on  the  fact  that  a  wing  with  a  practically  streamline  flow 
produces  a  series  of  whirling  vortices  (Whirlpools)  in  the 
wake  of  the  wings,  and  that  the  forward  movement  of  the 
plane  produces  the  energy  that  is  stored  in  the  vortices. 
The  relation  betw^een  these  vortices  is  such,  that  when 
their  motion  is  destroyed,  they  give  up  their  energy  and 
produce  a  lifting  reaction  by  their  downward  momentum. 
The  upw^ard  reaction  on  the  wing  is  thus  equal  and  oppo- 
site to  the  downward  momentum  of  the  air  vortices. 

Drag  Components.  At  large  angles  of  incidence  where 
turbulence  exists,  the  lift  and  also  the  drag  are  nearly 
proportional  to  the  velocity  squared  (V^).  Where  little 
turbulence  exists,  and  where  the  air  stream  hugs  the  sur- 
face closely,  the  drag  is  due  largely  to  skin  friction,  and 
consequently  this  part  of  the  drag  varies  according  to 


10-1  AEROFOILS 

Zahm's  law  of  friction  (V^-^^).  For  this  reason  it  is  difficult 
to  estimate  the  difference  in  drag  produced  by  differences 
in  velocity,  since  the  two  drag  components  vary  at  dif- 
ferent rates,  and  there  is  no  fixed  proportion  between 
them.  Since  the  frictional  drag  does  not  increase  in  pro- 
portion to  the  area,  but  as  A^-®^,  difficulty  is  also  experi- 
enced in  estimating  the  drag  of  a  full  size  wing  from  data 
furnished  by  model  tests. 

Incidence  and  Lift.  Up  to  the  burble  point  the  lift  in- 
creases with  an  increase  in  the  angle ;  but  not  at  a  uniform 
rate  for  any  one  aerofoil,  nor  at  the  same  rate  for  different 
aerofoils.  The  drag  also  increases  with  the  angle,  but 
more  rapidly  than  the  lift  after  an  incidence  of  about  4° 
is  passed,  hence  the  lift-drag  ratio  is  less  at  angles  greater 
than  4°.  Decreasing  the  angle  beloW  4°  also  decreases 
the  lift-drag,  but  not  so  rapidly  as  with  the  larger  angles. 
At  the  angle  of  *'No  Lift"  the  drag  is  principally  due  to 
skin  friction. 

Fig.  3  shows  a  typical  lift  and  incidence  chart  that  gives 
the  relation  between  the  angle  of  incidence  ^  and  the 
lift  coefficient.  This  curve  varies  greatly  for  different 
forms  of  aerofoils  both  in  shape  and  numerical  value,  and 
it  is  only  given  to  show  the  general  form  of  such  a  graph. 
The  curve  lying  to  the  left,  and  above  the  curve  for  the 
''Flat  plate/'  is  the  curve  for  the  particular  aerofoil  shown 
above  the  chart.  The  "Lift-Coefficients"  at  the  left  hand 
vertical  edge  correspond  to  the  coefficient  Ky,  although 
these  must  be  multiplied  by  a  factor  to  convert  them  into 
values  of  Ky.  As  shown,  they  are  in  terms  of  the  "Abso- 
lute units  used  by  the  National  Physical  Laboratory  and 
to  convert  them  into  the  Ky  unit  they  must  be  multiplied 
by  0.005 IV^  where  V  is  in  miles  per  hour,  or  0.00236v^ 
where  v  =  feet  per  second.  The  incidence  angle  is  in 
degrees. 

It  will  be  noted  that  the  lift  of  the  aerofoil  is  greater 
than  that  of  the  plate  at  every  angle  as  with  nearly  every 


AEROFOILS 


105 


practical  aerofoil.  The  aerofoil  has  a  lift  coefficient  of 
0.0025  at  the  negative  angle  of  — 3°,  while  the  lift  of  the 
fiat  plate  of  course  becomes  zero  at  0°.  As  the  incidence  of 
the  aerofoil  increases  the  Hft  coefficient  also  increases, 
until  it  reaches  a  maximum  at  the  burble  point  (Stalling 
angle)  of  about  11.5°.  An  increase  of  angle  from  this  point 


T 


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A 

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Z    5 

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Fig.   3.      Chart  Showing  Relation  Between  Incidence  and  Lift. 

causes  the  lift  coefficient  to  drop  rapidly  until  it  reaches 
a  minimum  lift  coefficient  of  0.46  at  17°.  The  fiat  plate  as 
shown,  reaches  a  maximum  at  the  same  angle,  but  the  lift 
of  the  plate  does  not  drop  off  as  rapidly.  The  maximum 
coefficient  of  the  aerofoil  is  0.58  and  of  the  plate  0.41.  The 
rapid  drop  in  pressure,  due  to  the  air  stream  breaking 
away  at  the  burble  point,  is  clearly  shown  by  the  sharp 
peak  in  the  aerofoil  curve.    The  sharpness  of  the  drop 


106  AEROFOILS 

varies  among  different  aerofoils,  the  peaks  in  some  forms 
being  very  flat  and  uniform  for  quite  a  distance  in  a  hori- 
zontal direction,  while  others  are  even  sharper  than  that 
shown.  Everything  else  being  equal,  an  aerofoil  with  a 
fiat  peak  is  the  more  desirable  as  the  lift  does  not  drop 
off  so  rapidly  in  cases  where  the  aviator  exceeds  the  crit- 
ical angle,  and  hence  the  tendency  to  stall  the  machine  is 
"lot  as  great.  This  form  of  chart  is  probably  the  simplest 
form  to  read.  It  contains  only  one  quantity,  the  lift-co- 
efficient, and  it  shows  the  small  variations  more  clearly 
than  other  types  of  graphs  in  which  the  values  of  Kx,  lift- 
drag,  and  the  resultant  force  are  all  given  on  a  single 
sheet. 

Center  of  Pressure  Movement.  As  in  the  case  of  the  flat 
plate  the  center  of  pressure  on  an  aerofoil  surface  varies 
with  the  angle  of  incidence,  but  unlike  the  plate  the  center 
of  pressure  (C.  P.)  moves  backward  with  a  decrease  in 
angle.  The  rapidity  of  travel  depends  upon  the  form  of 
aerofoil,  in  some  types  the  movement  is  very  great  with 
a  small  change  in  the  angle,  while  in  others  the  movement 
is  almost  negligible  through  a  wide  range.  In  general, 
aerofoils  are  inherently  unstable,  since  the  C.  P.  moves 
toward  the  trailing  edge  with  decreased  angles,  and  tends 
to  aggravate  a  deficiency  in  the  angle.  If  the  angle  is  too 
small,  the  backward  movement  tends  to  make  it  still 
smaller,  and  with  an  increasing  angle  the  forward  move- 
ment of  the  center  of  pressure  tends  to  make  the  angle 
still  greater. 

Fig.  4  is  a  diagram  showing  the  center  of  pressure 
movement  for  a  typical  aerofoil  with  the  aerofoil  at  the 
top  of  the  chart.  The  left  side  of  the  chart  represents  the 
leading  edge  of  the  aerofoil  and  the  right  side  is  the  trail- 
ing edge,  while  the  movement  in  percentages  of  the  chord 
length  is  shown  by  the  figures  along  the  lower  line.  Thus 
figure  ".3"  indicates  that  the  center  of  pressure  is  located 
0.3  of  the  chord  from  the  leading  edge.   In  practice  it  is 


AEROFOILS 


107 


usual  to  measure  the  distance  of  the  C.  P.  from  the  leading 
edge  in  this  way. 

For  an  example  in  the  use  of  the  chart,  let  us  find  the 
location  of  the  C.  P.  at  angles  of  0°,  3°  and  7°.  Starting 
with  the  column  of  degrees  at  the  left  hand  edge  of  the 
chart,  find  0°,  and  follow  along  the  dotted  line  to  the  right 

a 

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A 

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Chord. 

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i^/fOfUti  £OQ£. 

rpAa/A/QEDOt. 

Fig.  4.     Chart  Giving  Relation  Between  Incidence  and  C.P.  Movement. 


until  the  curve  is  reached.  From  this  point  follow  down 
to  the  lower  row  of  figures.  It  will  be  found  that  at  0° 
the  C.  P.  lies  about  half  way  between  0.5  and  0.6,  or  more 
exactly  at  0.55  of  the  chord  from  the  leading  edge.  Simi- 
larly at  3°  the  C.  P.  is  at  0.37  of  the  chord,  and  at  7°  is  at 
0.3  of  the  chord.  From  11°  to  19°,  the  C.  P.  for  this  par- 
ticular aerofoil  moves  very  little,  remaining  almost  con- 


108  AEROFOILS 

stant  at  0.25  of  the  chord.  Reducing  the  angle  from  3*^ 
causes  the  C.  P.  to  retreat  very  rapidly  to  the  rear,  so 
that  at  — 1°  the  C.  P.  is  at  0.8  of  the  chord,  or  very  near 
the  trailing  edge  of  the  wing. 

Other  Forms  of  Charts.  The  arrangement  of  wing  per- 
formance charts  differs  among  the  various  investigators. 
Some  charts  show  the  lift,  drag,  lift-drag  ratio,  angle  of 
incidence,  center  of  pressure  movement,  and  resultant 
pressure  on  a  single  curve.  This  is  very  convenient  for 
the  experienced  engineer,  but  is  somewhat  complicated 
for  the  beginner.  Whatever  the  form  of  chart,  there  should 
be  an  outline  drawing  of  the  aerofoil  described  in  the  chart. 

Fig.  5  shows  a  chart  of  the  "Polar"  variety  in  which 
four  of  the  factors  are  shown  by  a  single  curve.  This 
type  was  originated  by  Eiffel  and  is  generally  excellent, 
except  that  the  changes  at  small  angles  are  not  shown 
very  clearly  or  sharply.  The  curve  illustrates  the  proper- 
ties of  the  ''KaufTman"  wing,  or  better  known  as  the 
^'Eiffel  No.  7)7 r  A  more  complete  description  of  this 
aerofoil  will  be  found  under  the  chapter  "Practical  Wing 
Sections."  A  single  curve  is  marked  at  different  points 
with  the  angle  of  incidence  (0°  to  12°).  The  column  at 
the  left  gives  the  lift-coefficient  Ky,  while  the  row  at  the 
bottom  of  the  sheet  gives  the  drag-coefficients  Kx.  At  the 
top  of  the  chart  are  the  lift-drag  ratios,  each  figure  being 
at  the  end  of  a  diagonal  line.  In  this  way  the  lift,  drag,  lift- 
drag  and  angle  of  incidence  are  had  from  a  single  curve. 

Take  the  characteristics  at  an  angle  of  10  degrees  for 
example.  Find  the  angle  of  10°  on  the  curve,  and  follow 
horizontally  to  the  left  for  Ky.  The  lift-coefficient  will  be 
found  to  be  0.0026  in  terms  of  miles  per  hour  and  pounds 
per  square  foot.  Following  down  from  10°,  it  will  be 
found  that  the  drag-coefficient  Kx  =  0.00036.  Note  the 
diagonal  lines,  and  that  the  10°  point  lies  nearest  to  the 
diagonal  headed  7  at  the  top  of  the  chart.  (It  is  more 
nearly  a  Hft-drag  ratio  of  7.33  than  7.)    In  the  same  way 


AEROFOILS 


109 


it  will  be  found  that  an  angle  of  8  degrees  lies  almost  ex- 
actly on  the  lift-drag  diagonal  marked  9.  The  best  lift- 
drag  is  reached  at  about  2  degrees  at  which  point  it  is 
shown  as  17.0.  The  best  lift-coefficient  Ky  is  0.00276  at  12 
degrees. 

A  third  class  of  chart  is  shown  by  Fig.  6.   This  single 


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Fig.  5.     Polar  Type   Chart   Originated   by   Eiffel. 


chart  shows  three  of  the  factors  by  means  of  three  curves  ; 
one  for  the  lift-coefficient,  one  for  the  drag-coefficient, 
and  one  for  the  C.  P.  movement.  Follow  the  solid  curves 
only,  for  the  dotted  lines  are  for  comparison  with  the 
results  obtained  by  another  laboratory  in  checking  the 
characteristics  of  the  wing.  The  curves  refer  to  the  R.  A. 
F. — 6  section  described  in  the  chapter  on  ''Practical  Wing 


110 


AEROFOILS 


Sections."  The  lift-coefficients  K^  will  be  found  at  the 
right  of  the  chart  with  the  drag-coefficients  Kx  at  the  left 
and  in  the  lower  column  of  figures.  The  upper  column  at 
the  left  is  for  the  C.  P.  movement  and  gives  the  C.  P. 
location  in  terms  of  the  chord  length.  The  angles  of  in- 
cidence will  be  found  at  the  bottom.  Values  are  in  terms 
of  pounds  per  square  foot,  and  miles  per  hour. 

In  using  this  chart,  start  with  the  angle  of  incidence  at 


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Fig. 


■A'      -2?        O         £•       A'        6'        «•        10*       12'       I4-*       i6*      16' 

6.     Chart  of  R.A.F.-6  Wing  Section  with  Three  Independent  Curves. 


AEROFOILS  111 

the  bottom,  and  follow  up  vertically  to  the  lift  or  drag 
curves.  If  the  value  of  Ky  is  desired,  proceed  from  the 
required  incidence  and  up  to  the  "Lift"  curve,  then  hori- 
zontally to  the  right.  To  obtain  the  drag,  follow  up  from 
the  angle  of  incidence  to  the  "drift"  curve,  and  then  hori- 
zontally to  the  left.  For  the  position  of  the  C.  P.,  trace 
up  from  angle  until  the  "Center  of  Pressure"  curve  is 
reached,  and  then  across  horizontally  to  the  left.  If  the 
angle  of  8  degrees  is  assumed,  the  lift-coefficient  will  be 
found  as  Ky  =  0.0022,  the  drag  Kx  =  0.00016,  and  the 
center  of  pressure  will  be  located  at  0.32  of  the  chord 
from  the  leading  edge.  This  test  was  made  with  the  air 
density  at  0.07608  pounds  per  cubic  foot,  and  at  a  speed 
of  29.85  miles  per  hour.  The  peak  at  the  burble  point  is 
fairly  flat,  and  gives  a  good  range  of  angle  before  the  lift 
drops  to  a  serious  extent.  The  aerofoil  R.  A.  F. — 6  is  a 
practical  wing  form  used  in  many  machines,  and  this 
fact  should  make  the  chart  of  special  interest. 

Surface  Calculations.  The  calculation  of  lift  and  drag 
for  an  aerofoil  are  the  same  as  those  for  a  flat  plate,  that 
is,  the  total  lift  is  expressed  by  the  formula : 
L  =  KyAV-  where  A  is  the  area  in  square  feet,  and  V  is 
the  velocity  in  miles  per  hour.  From  this  primary  equa- 
tion, the  values  of  the  area  and  velocity  may  be  found  by 
transposition.  k_ 

A  =  L/KyV^  and  V  =  L/  KyA.         ^<  >    -    A  ^f  "■ 

The  drag  can  be  found  from  the  old  equation,  D  =  Kx 
AV-,  or  by  dividing  the  lift  by  the  lift-drag  ratio  as  in 
the  case  of  the  flat  plate. 

Example :  A  wing  of  the  R.  A.  F. — 6  form  has  an  area  of 
200  square  feet,  and  the  speed  is  60  miles  per  hour.  What 
is  the  lift  at  6°  incidence? 

Solution.  From  Chart  No.  6  the  lift  coefficient  Ky  is 
0.00185  at  6°,  hence  the  total  lift  is  :  L  =  KyAV^  =  0.00185 
x200x(60x60)  =  1332  pounds.  With  an  angle  of  8  degrees, 
and  with  the  same  speed  and  area,  the  Hft  becomes, 


112  AEROFOILS 

L=0.0022x200x  (60x60)  =  1584  pounds.  The  drag  co- 
efficient Kx  at  an  angle  of  6°  is  0.00012,  and  at  8°  is  0.00016. 
The  drag  at  6°  becomes  D=KxAV2  =  0.00012x200x(60x 
60)  =86.4  pounds.  The  lift-drag  ratio  at  this  angle  is 
L/D  =  1332/86.4  =  15.4.  The  drag  at  8°  is  D  =  K^AV^  = 
0.00016  X  200  X  (60  x  60)  =  115.2  pounds.  The  lift-drag 
at  8°  is  L/D  =  1584/115.2=  13.8. 

Forces  Acting  on  Aerofoil.  Fig.  7  is  a  section  through 
an  aerofoil  of  a  usual  type,  with  a  concave  under-surface 
In  an  aerofoil  of  this  character  all  measurements  are 
made  from  the  chordal  line  X-X  which  is  a  straight  line 
drawn  across  (and  touching)  the  entering  and  trailing 
edges  of  the  aerofoil.  The  angle  made  by  X-X  with  the 
horizontal  is  the  angle  of  incidence  (i).  The  width  of  the 
section,  measured  from  tip  to  tip  of  the  entering  and  trail- 
ing edges,  is  called  the  *'Chord."  In  this  figure  the  enter- 
ing edge  is  at  the  left.  The  direction  of  lift  is  "Up"  or  as 
in  the  case  of  any  aerofoil,  acts  away  from  the  convex 
side. 

In  the  position  shown,  with  horizontal  motion  toward 
the  left,  the  lift  force  is  indicated  by  L,  and  the  horizontal 
drag  force  by  D,  the  direction  of  their  action  being  indi- 
cated by  the  arrow  heads.  The  force  that  is  the  resultant 
of  the  lift  and  drag,  lies  between  them,  and  is  shown  by 
R.  The  point  at  which  the  line  of  the  resultant  force  in- 
tersects the  chordal  line  X-X  is  called  the  "Center  of 
Pressure"  (C.  P.)  The  resultant  is  not  always  at  right 
angles  to  the  chordal  line  as  shown,  but  may  lie  to  either 
side  of  this  right  angle  line  according  to  the  angle  of 
incidence  (i).  A  force  equal  to  and  in  the  same  direction 
as  R,  will  hold  the  forces  L  and  D  in  equilibrium  if 
applied  at  the  center  of  pressure  (C.  P.)  Owing  to  the 
difference  in  the  relative  values  of  L  and  D  at  various 
angles  of  incidence,  the  angle  made  by  R  with  the  chordal 
line  must  vary.  The  lift  and  drag  are  always  at  right 
angles  to  one  another.    The  resultant  can  be  found  by 


AEROFOILS 


113 


drawing  both  the  lines  L  and  D  through  the  C.  P.,  and 
at  right  angles  to  one  another,  and  then  closing  up  the 
parallelogram  by  drawing  lines  parallel  to  L  and  D  from 
the  extreme  ends  of  the  latter.  The  resultant  force  in 
direction  and  extent  will  be  the  diagonal  R  drawn  across 
the  corners  of  the  parallelogram. 

The  forces  acting  on  the  upper  and  lower  surfaces  are 
different,  both  in  direction  and  magnitude,  owing  to  the 
fact  that  the  upper  and  lower  surfaces  do  not  contribute 
equally  to  the  support  of  the  aerofoil.  The  upper  surface 


Fig.  7.     Forces  Acting  on  an  Aerofoil,  Lift,  Drag,  and  Resultant. 
Wind  Is  from  Left  to  Right. 


Relative 


contributes  from  60  to  80  per  cent  of  the  total  lift.  A 
change  in  the  outline  of  the  upper  curved  surface  vitally 
affects  both  the  lift  and  lift-drag,  but  a  change  in  the 
lower  surface  affects  the  performance  to  an  almost  negli- 
gible amount. 

In  the  case  of  thin  circular  arched  plates  the  curvature 
has  a  much  more  pronounced  effect.  When  the  curvature 
of  a  thin  plate  is  increased,  both  the  upper  and  lower 
surfaces  are  increased  in  curvature,  and  this  undoubtedly 
is  the  cause  of  the  great  increase  in  the  lift  of  the  sheet 
metal  aerofoils  tested  by  Eiffel. 

The    drag   component   of  the   front   upper   surface    is 


114  AEROFOILS 

^'Negative,"  that  is,  acts  with  the  horizontal  force  instead 
of  against  it.  The  lower  surface  drag  component  is  of 
course  opposed  to  the  horizontal  propelHng  force  by- 
enough  to  wholly  overcome  the  assisting  negative  drag 
force  of  the  front  upper  surface.  The  resultants  vary  from 
point  to  point  along  the  section  of  the  aerofoil  both  in 
extent  and  direction.  A  resultant  true  for  the  entering 
edge  would  be  entirely  different  at  a  point  near  the  trail- 
ing edge. 

Distribution  of  Pressure.  To  fully  understand  the  rela- 
tive pressures  and  forces  acting  on  different  parts  of  the 
aerofoil  we  must  refer  to  the  experimental  results  ob- 
tained by  the  Eiffel  and  the  N.  P.  L.  laboratories.  In 
these  tests  small  holes  were  drilled  over  the  aerofoil  sur- 
face at  given  intervals,  each  hole  in  turn  being  connected 
to  a  manometer  or  pressure  gauge,  and  the  pressure  at 
that  point  recorded.  While  the  reading  was  being  taken, 
the  wind  was  passed  over  the  surface  so  that  the  pressures 
corresponded  to  actual  working  conditions.  It  was  found 
that  the  pressure  not  only  varied  in  moving  from  the  en- 
tering to  trailing  edge,  but  that  it  also  varied  from  the 
center  to  the  tips  in  moving  along  the  length  of  the  plane. 
The  rate  of  variation  differed  among  different  aerofoils, 
and  with  the  same  aerofoil  at  different  angles  of  incidence. 

On  the  upper  surface,  the  suction  or  vacuum  was  gen- 
erally very  high  in  the  immediate  vicinity  of  the  entering 
edge.  From  this  point  it  decreased  until  sometimes  the 
pressure  was  actually  reversed  near  the  trailing  edge  and 
at  the  latter  point  there  was  actually  a  downward  pres- 
sure acting  against  the  lift.  The  positive  pressure  on 
the  under  surface  reached  a  maximum  more  nearly  at  the 
center,  and  in  many  cases  there  was  a  vacuum  near  the 
entering  edge  or  at  the  trailing  edge.  With  nearly  all 
aerofoils,  an  increase  in  the  curvature  resulted  in  a  de- 
cided increase  in  the  vacuum  on  the  upper  surface,  par- 
ticularly with  thin  aerofoils  curved  to  a  circular  arc. 


AEROFOILS 


115 


By  taking  the  sum  of  the  pressures  at  the  various  parts 
of  the  surface,  it  was  found  that  the  total  corresponded 
to  the  lift  of  the  entire  aerofoil,  thus  proving  the  correct- 
ness of  the  investigation.  The  sum  of  the  drag  forces 
measured  at  the  different  openings  gave  a  lower  total  than 
the  total  drag  measured  by  the  balance,  and  this  at  once 
suggests  that  the  difference  was  due  to  the  skin  friction 
eft'ect  that  of  course  gave  no  pressure  indication.  The 
truth  of  this  deduction  is  still  further  proved  by  the  fact 
that  the  drag  values  were   more  nearly  equal  at  large 


Fig.  8.     Pressure    Distribution    for   Thin    Circular    Section, 
the  Effect  of  Increasing  the  Camber.      (Eiffel) 


Fig.    9.    Shows 


angles  where  the  turbulence  formed  a  greater  percentage 
of  the  total  drag. 

Figs.  8,  9,  10,  11,  12  are  pressure  distribution  curves 
taken  along  the  section  of  several  aerofoil  surfaces.  These 
are  due  to  Eiffel.  In  Fig.  8  is  the  pressure  curve  for  a 
thin  circular  aerofoil  section,  the  depth  of  the  curve  meas- 
ured from  the  chordal  line  being  1/13.5  of  the  chord.  The 
vacuum  distribution  of  the  upper  surface  is  indicated  by 
the  upper  dotted  curve,  while  the  pressure  on  the  bottom 
surface  is  given  by  the  solid  curve  under  the  aerofoil. 


116  AEROFOILS 

The  pressures  are  given  by  the  vertical  column  of  figures 
at  the  right  and  are  in  terms  of  inches  of  water,  that  is, 
the  pressure  required  for  the  support  of  a  w^ater  column 
of  the  specified  height.  Figures  lying  above  0  and  marked 
( — ),  refer  to  a  vacuum  or  negative  pressures,  while  the 
figures  below  zero  are  positive  pressures  above  the  atmos- 
pheric. The  entering  edge  is  at  the  right,  and  the  angle 
of  incidence  in  all  cases  is  6°. 

It  will  be  seen  that  the  vacuum  jumps  up  very  suddenly 
to  a  maximum  at  the  leading  edge,  and  again  drops  as 
suddenly  to  about  one-half  the  maximum.  From  this 
point  it  again  gradually  increases  near  the  center,  and 
then  declines  toward  the  trailing  edge.  It  will  also  be 
seen  that  the  pressure  on  the  lower  surface,  given  by  the 
solid  curve,  is  far  less  than  the  pressure  due  to  the  upper 
surface.  Since  the  lower  pressure  curve  crosses  up,  and 
over  the  zero  line  at  a  point  near  the  trailing  edge,  it  is 
evident  that  the  supper  surface  near  the  trailing  edge  is 
under  a  positive  pressure,  or  a  pressure  that  acts  down  and 
against  the  lift.  The  pressures  in  any  case  are  very  minute, 
the  maximum  suction  being  0.3546  inch  of  water,  while  the 
maximum  pressure  on  the  under  surface  is  only  0.085 
inch. 

Fig.  9  shows  the  effect  of  increasing  the  curvature  or 
camber,  the  aerofoil  in  this  case  having  a  depth  equal  to 
1/7  the  chord,  or  nearly  double  the  camber  of  the  first. 
The  sharp  peak  at  the  entering  edge  of  the  pressure  curve 
is  sHghtly  reduced,  but  the  remaining  suction  pressures 
over  the  rest  of  the  surface  are  much  increased,  indicating 
a  marked  increase  in  the  total  pressure.  The  pressure 
at  the  center  is  now  nearly  equal  to  the  front  peak,  and 
the  pressure  is  generally  more  evenly  distributed.  There 
is  a  vacuum  over  the  entire  upper  surface  and  a  positive 
pressure  over  the  lower.  The  general  increase  in  pressure 
due  to  the  increased  camber  is  the  result  of  the  greater 
downward   deviation  of  the   air  stream,   and   the   corre- 


AEROFOILS  117 

spending  greater  change  in  the  momentum  of  the  air. 
The  speed  at  which  the  tests  were  made  was  10  meters 
per  second,  or  22.4  miles  per  hour.  The  curves  are  only 
true  at  the  center  of  the  aerofoil  length  and  for  an  aspect 
ratio  of  6. 

The  average  pressure  over  the  entire  surface  in  Fig. 
8  is  1.202  pounds  per  square  foot,  and  that  of  Fig.  9  is 
1.440  pounds,  a  difference  of  0.238  pound  per  square  foot 
due  to  the  doubling  camber  (16.5  per  cent).  Another  aero- 
foil with  a  camber  of  only  1/27  gave  an  average  pressure 
of  0.853  pound  per  square  foot  under  the  same  conditions. 
A  flat  plane  gave  0.546.  Tabulation  of  these  values  will 
show  the  results  more  clearly. 

Camber      Av.  Pres.      Inc.  in  Pres.  Efficiency 

Of  Surface  Per  Sq.  Ft.  in  Lbs./Sq.  Ft.   Top.         Bottom. 

Flat  Plane.     0.546  0.000  0.89  0.11 

1/27  0.853  0.307  0.72  0.28 

1/13.5       1.202  0.349  0.71  0.29 

1/7  1.440  0.238  0.59  0.41 

In  this  table,  the  "Efficiencies"  are  the  relative  lift 
efficiencies  of  the  top  and  bottom  surfaces.  For  example, 
in  the  case  of  the  1/7  camber  the  top  surface  lifts  59  per 
cent,  and  the  bottom  41  per  cent  of  the  total  lift. 

Fig.  10  is  a  thin  aerofoil  of  parabolic  form,  while  Fig. 
11  is  an  approximation  to  the  compafaiivcly  thick  wi?Tg  oT 
a  bird.  In  both  these  sections  it  will  be  noted  that  the 
front  peak  is  not  much  greater  than  the  secondary  peak, 
and  that  the  latter  is  nearer  the  leading  edge  than  with 
the  circular  aerofoils.  Also  that  the  drop  between  the 
peaks  is  small  or  entirely  lacking.  The  lower  surface  of 
the  trailing  edge  is  subjected  to  a  greater  down  pressure 
in  the  case  of  the  thin  parabola,  and  there  is  also  a  con- 
siderable down  pressure  on  the  upper  leading  edge.  The 
pressure  in  Fig.  10  is  1.00  pound  per  square  foot,  and  that 


118 


AEROFOILS 


of  No.  11  is  1.205,  while  the  efficiency  of  the  top  surfaces 
is  respectively  72  and  74  per  cent. 

Fig.  12  shows  the  effect  of  changing  the  angle  of  the 
bird  wing  from  zero  to  8  degrees.  The  lift  per  square  foot 
in  each  case  is  shown  at  the  upper  left  hand  corner  of  the 
diagram  while  the  percentages  of  the  upper  and  lower 
surface  lifts  are  included  above  and  below  the  wing.  For 
these  curves  I  am  indebted  to  E.  R.  Armstrong,  formerly 


z^^zzji^r 


Fig.   10.     Thin    Parabolic    Aerofoil    with    Pressure    Distribution.       Fig.     11. 
Pressure  Distribution  of  Thick  Bird's  Wing  Type.      (Eiffel) 


of  "Aero  and  Hydro."  As  the  angle  is  increased,  the 
suction  of  the  upper  surface  is  much  increased  (0.541  to 
1.370  pounds  per  square  foot),  and  the  pressure  at  the 
leading  edge  increases  from  depression  to  a  very  long 
thin  peak.  The  maximum  under  pressure  is  not  much 
increased  by  the  angle,  but  its  distribution  and  average 
pressure  are  much  altered.  At  0°  and  2°  the  usual  pres- 
sure is  reduced  to  a  vacuum  over  the  front  of  the  section 
as  shown  by  the  lower  curve  crossing  over  the  upper  side 
of  the  wing,  and  at  this  point  the  under  surface  sucks 
down  and  acts  against  the  lift. 


AEROFOILS 


119 


Distribution  of  Drag  Forces.  The  drag  as  well  as  the  lift 
changes  in  both  direction  and  magnitude  for  different 
points  on  the  wing.  In  the  front  and  upper  portions  the 
drag  is  "Negative,"  that  is,  instead  of  producing  head  re- 
sistance to  motion  it  really  acts  with  the  propelling  force. 
Hence  on  the  upper  and  front  portions  the  lift  is  ob- 
tained with  no  expenditure  of  power,  and  in  fact  thrust  is 
given  up  and  added  to  that  of  the  propeller.  The  remain- 
ing drag  elements  at  the  rear,  and  on  the  lower  surface, 
of  course  more  than  overcome  this  desirable  tendency 


•BtR?' 


Fig,   12.     Effect   of   Incidence   Changes   on   the   Pressure   Distribution   of  a 
Thick  Bird's  Wing.      (After  Eiffel) 


and  give  a  positive  drag  for  the  total  wing.  The  distribu- 
tion is  shown  by  Fig  13  which  gives  the  lift,  drag  and 
resultant  forces  at  a  number  of  different  points  on  two 
circular  arc  aerofoils  having  cambers  of  1/13.5  and  1/7 
respectively.  In  this  figure,  the  horizontal  drag  forces  are 
marked  D  and  d,  and  the  direction  of  the  drag  is  shown 
by  the  arrows.  The  lift  is  shown  by  L  and  the  resultant 
by  R  as  in  the  Fig.  7. 

As  shown,  the  arrows  pointing  to  the  right  are  the 
"Negative"  drag  (d)  forces  that  assist  in  moving  the 
plane  forward,  while  the  drag  indicated  by  arrows  (D) 
pointing  to  the  left  are  the  drag  forces  that  oppose  or 


120 


AEROFOILS 


resist  the  horizontal  motion.  With  the  smaller  camber 
(1/13.5)  the  drag  forces  are  very  much  smaller  than  those 
with  the  heavier  camber  of  1/7,  and  the  negative  drifts  (d) 
are  correspondingly  smaller.  All  of  the  drag  due  to  the 
lower  surface,  point  to  the  left  (D),  and  hence  produce 
head  resistance  to  flight.  The  drag  to  the  rear  of  the 
center  of  the  upper  surface  are  the  same.  In  front  of 
the  upper  center  we  have  right  hand,  or  negative  drifts 
(d),  that  aid  the  motion.  These  forward  forces  obtained 
by  experiment  prove  the  correctness  of  Lilienthal's  ''For- 
ward Tangential"  theory  advanced  many  years  ago. 


Fig.  13. 


Direction    of    Drag    Over    Different    Portions    of    Circular    Arc 
Aerofoils. 


Distribution  on  Practical  Wings.  With  the  exception 
of  the  bird  wing,  the  distributions  have  been  given  for 
thin  plates  that  are  of  little  value  on  an  aeroplane.  They 
do  not  permit  of  strong  structural  members  for  carry- 
ing the  load.  The  actual  wing  must  have  considerable 
thickness,  as  shown  by  the  aerofoils  in  Figs.  1,  2,  3,  etc., 
and  are  of  approximately  stream  line  form.  Fig.  14  shows 
the  distribution  for  actual  aeroplane  wings :  (a)  Wright, 
(b)  M.  Farman,  (c)  Breguet,  (d)  Bleriot  11. (d),  (e)  Bler- 
iot  11-bis.  The  Wright  wing  is  very  blunt  and  has  an  ex- 
ceedingly high  lift  at  the  leading  edge.  The  M.  Farman, 
which  is  slightly  less  blunt,  has  a  similar  but  lower  front 
peak.  The  Breguet  is  of  a  more  modern  type  with  the 
maximum  thickness  about  25  per  cent  from  the  leading 


AEROFOILS 


121 


edge.  The  latter  shows  a  remarkably  even  distribution 
of  pressure,  and  is  therefore  a  better  type  as  will  be  seen 
from  the  relative  lifts  of  0.916  and  0.986  pounds  per  square 
foot.  The  lift-drag  ratio  of  the  Breguet  is  also  better, 
owing  to  the  greater  predominance  of  the  negative  drag 


wR»rr(cL) 


.  PARKAN  (^       ,BRE<iUET  (0 


(cL)    (f)  6LCRK)T  r;^ 


Fig.   14a.     Distribution    for    Wright    Wing.       (b)    M.    Farman    Wing.       (c) 
Breguet  Wing,      (d)   Bleroit  Wing,      (e)   Bleriot  11-Bis. 


Fig.   15.     Pressure   Distribution   at   Various   Points   Along  the   Length    of  a 
Nieuport  Monoplane  Wing. 


components.  Decreasing  the  thickness  and  the  under- 
camber  of  Bleriot  11,  resulted  in  an  unusual  increase  of 
10  per  cent  of  the  under  pressure,  and  a  decrease  in  the 
suction,  shown  by  Bleriot  11-bis.  The  Bleriot  has  the 
sharpest  entering  edge  and  the  least  upper  pressure. 

In  the  above  practical  wing  sections  the  aspect  ratio 
is  variable,  being  the  same  in  the  test  model  as  in  the 
full-size  machine.     The  Bleriot  being  a  monoplane  has 


122 


AEROFOILS 


a  lower  aspect  ratio  (5),  than  the  biplanes  (a),  (b)  and 
(c).  The  Breguet  with  an  aspect  of  8  has  a  lift  of  0.986 
pounds  per  square  foot  as  against  the  0.781  of  the  Bleriot, 
and  undoubtedly  part  of  this  difference  is  due  to  the 
aspect  ratio.  The  pressure  falls  off  around  the  tips  as 
shown  by  the  successive  sections  taken  through  a  Nieu- 


LEAOG 

eoee 


i  1  i  1  1  !  1  1 


ILEADGI 
i  EDGE 


Pressure:    '^  Distribution 


PRESS.  =  O.QO" WATER,    lu 
PRESS.  =  0.65"  WATER,     g 

z 


RPgESS.-  0. 12  WATER 


P>RE5S.=  O.OSWATgg 


PRfeSS. »  0. 35"  WATER ■    {Jjl  PRE55.  =  "Q.OS WATER  ^^g 


Fig.  16.  Showing  Pressure  Distribution  on  the  Plan  View  of  a  Typical 
Wing,  Leading  Edge  Along  A-A,  Trailing  Edge  D-C.  Center  of 
Pressure  Marked  "C.P."  The  Proportion  Pressures  Are  Indicated 
by  the  Shading  on  the  Surface,  the  Pressure  Being  Negative  at  the 
Tips  and  Near  the  Rear  Edge. 

port  monoplane  wing  in  Fig.  15.  Section  (f)  was  taken 
near  the  body  and  shows  the  greater  lift.  Section  (g)  is 
midway  between  the  tips  and  body,  and  (h)  and  (i)  are 
progressively  nearer  the  tips.  As  we  proceed  toward  the 
tips  from  the  body  the  pressure  falls  off  as  shown  in  the 
sections,  this  reducing  from  1.07  to  0.55  pounds  per  square 
foot.  This  wing  also  thins  down  toward  the  tips  or 
''washes  out,"  as  it  is  called. 


CHAPTER  VI. 
PRACTICAL  WING  SECTIONS. 

Development  of  Modern  Wings.  The  first  practical 
results  obtained  by  Wright  Brothers,  Montgomery,  Cha- 
nute,  Henson,  Curtiss,  Langley,  and  others,  were  obtained 
by  the  use  of  cambered  wings.  The  low  value  of  the  lift- 
drag  ratio,  due  to  the  flat  planes  used  by  the  earlier  experi- 
menters, was  principally  the  cause  of  their  failure  to  fly. 
The  Wrights  chose  wings  of  very  heavy  camber  so  that 
a  maximum  lift  could  be  obtained  with  a  minimum  speed. 
These  early  wings  had  the  very  fair  lift-drag  ratio  of  12 
to  1.  Modern  wing  sections  have  been  developed  that 
give  a  lift-drag  ratio  of  well  over  20  to  1,  although  this  is 
attended  by  a  considerable  loss  in  the  lift. 

As  before  explained,  the  total  lift  of  a  wing  surface 
depends  on  the  form  of  the  wing,  its  area,  and  the  speed 
upon  which  it  moves  in  relation  to  the  air.  Traveling  at 
a  low  speed  requires  either  a  wing  with  a  high  lift  co-efii- 
cient  or  an  increased  area.  With  a  constant  value  for  the 
lift-drag  ratio,  an  increase  in  the  lift  value  of  the  wing 
section  is  preferable  to  an  increase  in  area,  since  the  larger 
area  necessitates  heavier  structural  members,  more 
exposed  bracing,  and  hence,  more  head  resistance.  Unfor- 
tunately, it  is  not  always  possible  to  use  the  sections  giv- 
ing the  heaviest  lift,  for  the  reason  that  such  sections 
usually  have  a  poor  lift-drag  ratio.  In  the  practical 
machine,  a  compromise  must  be  effected  between  the  drag 
of  the  wings  and  the  drag  or  head  resistance  of  the  struc- 
tural parts  so  that  the  combined  or  total  head  resistance 
will  be  at  a  minimum.     In  making  such  a  compromise, 

123 


124  WING  SECTIONS 

it  must  be  remembered  that  the  head  resistance  of  the 
structural  parts  predominates  at  high  speeds,  while  the 
drag  of  the  wings  is  the  most  important  at  low  speeds. 

In  the  early  days  of  flying,  the  fact  that  an  aeroplane 
left  the  ground  was  a  sufficient  proof  of  its  excellence, 
but  nowadays  the  question  of  efBciency  under  different 
conditions  of  flight  (performance)  is  an  essential.  Each 
new  aeroplane  is  carefully  tested  for  speed,  rate  of  climb, 
and  loading.  Speed  range,  or  the  relation  between  the 
lowest  and  highest  possible  flight  speeds,  is  also  of  increas- 
ing importance,  the  most  careful  calculations  being  made 
to  obtain  this  desirable  quality. 

Performance.  To  improve  the  performance  of  an  aero- 
plane, the  designer  must  increase  the  ratio  of  the  horse- 
power to  the  weight,  or  in  other  words,  must  either  use 
greater  horsepower  or  decrease  the  weight  carried  by  a 
given  power.  This  result  may  be  obtained  by  improve- 
ments in  the  motor,  or  by  improvements  in  the  machine 
itself.  Improvements  in  the  aeroplane  may  be  attained 
in  several  ways:  (1)  by  cutting  down  the  structural 
weight;  (2)  by  increasing  the  efficiency  of  the  lifting  sur- 
faces ;  (3)  by  decreasing  the  head  resistance  of  the  body 
and  exposed  structural  parts,  and  (4)  by  adjustment  of  the 
area  or  camber  of  the  wings  so  that  the  angle  of  incidence 
can  be  maintained  at  the  point  of  greatest  plane  efftciency. 
At  present  we  are  principally  concerned  with  item  (2), 
although  (4)  follows  as  a  directly  related  item. 

Improvement  in  the  wing  characteristics  is  principally 
a  subject  for  the  wind  tunnel  experimentalist,  since  with 
our  present  knowledge,  it  is  impossible  to  compute  the 
performance  of  a  wing  by  direct  mathematical  methods. 
Having  obtained  the  characteristics  of  a  number  of  wing 
sections  from  the  aerodynamic  laboratory,  the  designer  is 
in  a  position  to  proceed  with  the  calculation  of  the  areas, 
power,  etc.  At  present  this  is  rather  a  matter  of  elimina- 
tion, or  "survival  of  the  fittest,"  as  each  wing  is  taken 


WING  SECTIONS  125 

separately  and  computed  through  a  certain  range  of  per- 
formance. 

Wing  Loading.  The  basic  unit  for  wing  lift  is  the  load 
carried  per  unit  of  area.  In  English  units  this  is  expressed 
as  being  the  weight  in  pounds  carried  by  a  square  foot 
of  the  lifting  surface.  Practically,  this  value  is  obtained 
by  dividing  the  total  loaded  weight  of  the  machine  by 
the  wing  area.  Thus,  if  the  weight  of  a  machine  is  2,500 
pounds  (loaded),  and  the  area  is  500  square  feet,  the  "unit 
loading"  will  be:  w«=2,500/500  =  5  pounds  per  square 
foot.  In  the  metric  system  the  unit  loading  is  given  in 
terms  of  kilogrammes  per  square  meter.  Conversely, 
with  the  total  weight  and  loading  known,  the  area  can 
be  computed  by  dividing  the  weight  by  the  unit  loading. 
The  unit  loading  adopted  for  a  given  machine  depends 
upon  the  type  of  machine,  its  speed,  and  the  wing  sec- 
tion adopted,  this  quantity  varying  from  3.5  to  10  pounds 
per  square  foot  in  usual  practice.  As  will  be  seen,  the 
loading  is  higher  for  small  fast  machines  than  for  the 
slower  and  larger  types. 

A  very  good  series  of  wings  has  been  developed,  rang- 
ing from  the  low  resistance  type  carrying  5  pounds  per 
square  foot  at  45  miles  per  hour,  to  the  high  lift  wing, 
which  gives  a  lift  of  7.5  pounds  per  square  foot  at  the 
same  speed.  The  medium  lift  wing  will  be  assumed  to 
carry  6  pounds  per  square  foot  at  45  miles  per  hour.  The 
wing  carrying  7.5  pounds  per  square  foot  gives  a  great 
saving  in  area  over  the  low  lift  type  at  5  pounds  per 
square  foot,  and  therefore  a  great  saving  in  weight.  The 
weight  saved  is  not  due  to  the  saving  in  area  alone,  but  is 
also  due  to  the  reduction  in  stress  and  the  corresponding 
reduction  in  the  size  and  weight  of  the  structural  mem- 
bers. Further,  the  smaller  area  requires  a  smaller  tail 
surface  and  a  shorter  body.  A  rough  approximation  gives 
a  saving  of  1.5  pounds  per  square  foot  in  favor  of  the 
7.5  pound  wing  loading.     This  materially  increases  the 


126  WING  SECTIONS 

horsepower  weight  ratio  in  favor  of  the  high  lift  wing, 
and  with  the  reduction  in  area  and  weight  comes  an  im- 
provement in  the  vision  range  of  the  pilot  and  an  increased 
ease  in  handling  (except  in  dives).  The  high  lift  types 
in  a  dive  have  a  low  limiting  speed. 

As  an  offset  to  these  advantages,  the  drag  of  the  high 
lift  type  of  wing  is  so  great  at  small  angles  that  as  soon 
as  the  weight  per  horsepower  is  increased  beyond  18 
pounds  we  find  that  the  speed  range  of  the  low  resistance 
type  increases  far  beyond  that  of  the  high  lift  wing. 
According  to  Wing  Commander  Seddon,  of  the  English 
Navy,  a  scout  plane  of  the  future  equipped  with  low 
resistance  wings  will  have  a  speed  range  of  from  50  to 
150  miles  per  hour.  The  same  machine  equipped  with 
high  lift  wings  would  have  a  range  of  only  50  to  100  miles 
per  hour.  An  excess  of  power  is  of  value  with  low  resist- 
ance wings,  but  is  increasingly  wasteful  as  the  lift  co-effi- 
cient is  increased. 

Landing  speeds  have  a  great  influence  on  the  type  of 
wing  and  the  area,  since  the  low  speeds  necessary  for  the 
average  machines  require  a  high  lift  wing,  great  area,  or 
both.  With  the  present  wing  sections,  low  flight  speeds 
are  obtained  with  a  sacrifice  in  the  high  speed  values.  In 
the  same  way,  high  speed  machines  must  land  at  dan- 
gerously high  speeds.  At  present,  the  best  range  that 
we  can  hope  for  with  fixed  areas  is  about  two  to  one ; 
that  is,  the  high  speed  is  not  much  more  than  twice  the 
lowest  speed.  A  machine  with  a  low  speed  of  45  miles 
per  hour  cannot  be  depended  upon  to  safely  develop  a 
maximum  speed  of  much  over  90  miles  per  hour,  for  at 
higher  speeds  the  angle  of  incidence  will  be  so  diminished 
as  to  come  dangerously  near  to  the  position  of  no  lift.  In 
any  case,  the  travel  of  the  center  of  pressure  will  be  so 
great  at  extreme  wing  angles  as  to  cause  considerable 
manipulation  of  the  elevator  surface,  resulting  in  a. 
further  increase  in  the  resistance. 


WING  SECTIONS  127 

Resistance  and  Power.  The  horizontal  drag  (resist- 
ance) of  a  wing,  determines  the  power  required  for  its 
support  since  this  is  the  force  that  must  be  overcome  by 
the  thrust  of  the  propeller.  The  drag  is  a  component  of 
the  weight  supported  and  therefore  depends  upon  the 
loading  and  upon  the  efficiency  of  the  wing.  The  drag  of 
the  average  modern  wing,  structural  resistance  neglected, 
is  about  1/16  of  the  weight  supported,  although  there  are 
several  sections  that  give  a  drag  as  low  as  1/23  of  the 
weight.  The  denominators  of  these  fractions,  such  as 
*'16"  and  *'23,"  are  the  lift-drag  ratios  of  the  wing  sec- 
tions. 

Drag  in  any  wing  section  is  a  variable  quantity,  the  drag 
varying  with  the  angle  of  incidence.  In  general,  the  drag 
is  at  a  minimum  at  an  angle  of  about  4  degrees,  the 
value  increasing  rapidly  on  a  further  increase  or  decrease 
in  the  angle.  Usually  a  high  lift  section  has  a  greater 
drag  than  the  low  lift  type  at  small  angles,  and  a  smaller 
drag  at  large  angles,  although  this  latter  is  not  invariably 
the  case. 

Power  Requirements.  Power  is  the  rate  of  doing  work, 
or  the  rate  at  which  resistance  is  overcome.  With  a  con- 
stant resistance  the  power  will  be  increased  by  an  increase 
in  the  speed.  With  a  constant  speed,  the  power  will  be 
increased  by  an  increase  in  the  resistance.  Numerically, 
the  power  is  the  product  of  the  force  and  the  velocity  in 
feet  per  second,  feet  per  minute,  miles  per  hour,  or  meters 
per  second.  The  most  common  English  power  unit  is  the 
"horsepower,"  which  is  obtained  by  multiplying  the  resist- 
ing force  in  pounds  by  the  velocity  in  feet  per  minute, 
this  product  being  divided  by  33,000.  If  D  is  the  hori- 
zontal drag  in  pounds,  and  v  =  velocity  of  the  wing  in 
feet  per  minute,  the  horsepower  H  will  be  expressed  by — 

Dv 

H  = 

33,000 


128  WING  SECTIONS 

Since  the  speed  of  an  aeroplane  is  seldom  given  in  feet 
per  minute,  the  formula  for  horsepower  can  be  given  in 
terms  of  miles  per  hour  by — 

DV' 
H= 

375 
Where  V  =  velocity  in  miles  per  hour,  D  and  H  remain- 
ing as  before.  The  total  power  for  the  entire  machine 
would  involve  the  sum  of  the  wing  and  structural  drags, 
with  D  equal  to  the  total  resistance  of  the  machine. 
\  Example.  The  total  weight  of  an  aeroplane  is  found  to 
be  3,000  pounds.  The  lift-drag  ratio  of  the  wings  is  15.00, 
and  the  speed  is  90  miles  per  hour.  Find  the  power 
required  for  the  wings  alone. 

Solution.  The  total  drag  of  the  wings  wall  be :  D  = 
3,000/15  =  200  pounds.  The  horsepower  required:  H  = 
DV/375=200  X  90/375  =  48  horsepower.  It  should  be 
remembered  that  this  is  the  power  absorbed  by  the  wings, 
the  actual  motor  power  being  considerably  greater  owing 
to  losses  in  the  propeller.  With  a  propeller  efficiency  of 
70  per  cent,  the  actual  motor  power  will  become :  Hm  = 
48/0.70  =  68.57  for  the  wings  alone.  To  include  the  effi- 
ciency into  our  formula,  we  have — 

DV 

H  = 

375E 
where  E  =  propeller  efficiency  expressed  as  a  decimal. 
The  greater  the  propeller  efficiency,  the  less  will  be  the 
actual  motor  power,  hence  the  great  necessity  for  an 
efficient  propeller,  especially  in  the  case  of  pusher  type 
aeroplanes  where  the  wings  do  not  gain  by  the  increased 
slip  stream. 

The  propeller  thrust  must  be  equal  and  opposite  to  the 
drag  at  the  various  speeds,  and  hence  the  thrust  varies 
with  the  plane  loading,  wing  section,  and  angle  of  inci- 
dence.    Portions  of  the  wing  surfaces  that  lie  in  the  pro- 


WING  SECTIONS  129 

peller  slip  stream  have  a  greater  lift  than  those  lying 
outside  of  this  zone  because  of  the  greater  velocity  of  the 
slip  stream.  For  accurate  results,  the  area  in  the  slip 
stream  should  be  determined  and  calculated  for  the 
increased  velocity. 

Oftentimes  it  is  desirable  to  obtain  the  **Unit  drag"; 
that  is,  the  drag  per  square  foot  of  lifting  surface.  This 
can  be  obtained  by  dividing  the  lift  per  square  foot  by  the 
lift-drag  ratio,  care  being  taken  to  note  the  angle  at 
which  the  unit  drag  is  required. 

Advantages  of  Cambered  Sections  Summarized.  Mod- 
ern wing  sections  are  always  of  the  cambered,  double- 
surface  type  for  the  following  reasons : 

1.  They  give  a  better  lift-drag  ratio  than  the  fiat  sur- 

face, and  therefore  are  more  economical  in  the 
use  of  power. 

2.  In  the  majority  of  cases  they  give  a  better  lift  per 

square  foot  of  surface  than  the  fiat  plate  and 
require  less  area. 

3.  The  cambered  wings  can  be  made  thicker  and  will 

accommodate  heavier  spars  and  structural  mem- 
bers without  excessive  head  resistance. 

Properties  of  Modern  Wings.  The  curvature  of  a  wing 
surface  can  best  be  seen  by  cutting  out  a  section  along 
a  line  perpendicular  to  the  length  of  the  wing,  and  then 
viewing  the  cut  portion  from  the  end.  It  is  from  this 
method  of  illustration  that  the  different  wing  curves,  or 
types  of  wings,  are  known  as  'Sving  sections."  In  all 
modern  wings  the  top  surface  is  well  curved,  and  in  the 
majority  of  cases  the  bottom  surface  is  also  given  a 
curvature,  although  this  is  very  small  in  many  instances. 

Fig.  1.  shows  a  typical  wing  section  with  the  names 
of  the  different  parts  and  the  methods  of  dimensioning 
the  curves.  All  measurements  to  the  top  and  bottom 
surfaces  are  taken  from  the  straight  ''chordal   line"  or 


'/ 


130  WING  SECTIONS 

''datum  line"  marked  X-X.  This  line  is  drawn  across 
the  concave  undersurface  in  such  a  way  as  to  touch  the 
surface  only  at  two  points,  one  at  the  front  and  one  at 
the  rear  of  the  wing  section.  The  inclination  of  the  wing 
with  the  direction  of  flight  is  always  given  as  the  angle 
made  by  the  line  X-X  with  the  wind.  Thus,  if  a  certain 
wing  is  said  to  have  an  angle  of  incidence  equal  to  4 
degrees,  we  know  that  the  chordal  line  X-X  makes  an 
angle  of  4  degrees  with  the  direction  of  travel.  This  angle 
is  generally  designated  by  the  letter  (i),  and  is  also 
known  as  the  "angle  of  attack."  The  distance  from  the 
extreme  front  to  the  extreme  rear  edge  (width  of  wing) 
is  called  the  "chord  width"  or  more  commonly  "the  chord." 

In  measuring  the  curve,  the  datum  line  X-X  is  divided 
into  a  number  of  equal  parts,  usually  10,  and  the  lines 
1-2-3-4-5-6-7-8-9-10-11  are  drawn  perpendicular  to  X-X. 
Each  of  the  vertical  numbered  lines  is  called  a  "station," 
the  line  No.  3  being  called  "Station  3,"  and  so  on.  The 
vertical  distance  measured  from  X-X  to  either  of  the 
curves  along  one  of  the  station  lines  is  known  as  the 
"ordinate"  of  the  curve  at  that  point.  Thus,  if  we  know 
the  ordinates  at  each  station,  it  is  a  simple  matter  to  draw 
the  straight  line  X-X,  divide  it  into  10  parts,  and  then 
lay  off  the  heights  of  the  ordinates  at  the  various  sta- 
tions. The  distances  from  datum  to  the  upper  curve  are 
known  as  the  "Upper  ordinates,"  while  the  same  measure- 
ments to  the  under  surface  are  known  as  the  "Lower 
ordinates,"  This  method  allows  us  to  quickly  draw  any 
wing  section  from  a  table  that  gives  the  upper  and  lower 
ordinates  at  the  different  stations. 

A  common  method  of  expressing  the  value  of  the  depth 
of  a  wing  section  in  terms  of  the  chord  width  is  to  give 
the  "Camber,"  which  is  numerically  the  result  obtained 
by  dividing  the  depth  of  the  wing  curve  at  any  point  by 
the  width  of  the  chord.  Usually  the  camber  given  for 
a  wing  is  taken  to  be  the  maximum  camber;  that  is,  the 


WING  SECTIONS 


131 


camber  taken  at  the  point  of  greatest  depth.  Thus,  if  we 
hear  that  a  certain  wing  has  a  camber  of  0.089,  we  take 
it  for  granted  that  this  is  the  camber  at  the  deepest  por- 
tion of  the  wing.  The  correct  method  would  be  to  give 
0.089  as  the  "maximum  camber"  in  order  to  avoid  con- 
fusion. To  obtain  the  maximum  camber,  divide  the 
maximum  ordinate  by  the  chord. 

Example.     The  maximum  ordinate  of  a  certain  wing  is 
5  inches,  and  the  chord  is  40  inches.    What  is  the  max- 


5-^      4^ 

MAY. 


-CHORD  =C- 


STATION  NUMBERS. 
S  <2>  -7 


/o 


iLilA^ili.^ 


2/  =  LOV^ER  ORDINA TES 


CHOROAL  OR  DATUM  LINE. 

XJ  =  UPPER  OR  DIN  AT E5  • 


Fig.  1.  Section  Through  a  Typical  Aerofoil  or  Wing,  the  Parts  and  Measure- 
ments Being  Marked  on  the  Section.  The  Horizontal  Width  or 
"Chord"  Is  Divided  Into  10  Equal  Parts  or  "Stations,"  and  the 
Height  of  the  Top  and  Bottom  Curves  Are  Measured  from  the 
Chordal  Line  X-X  at  Each  Station.  The  Vertical  Distance  from  the 
Chordal  Line  Is  the  "Ordinate"  at  the  Point  of  Measurement. 


imum  camber?  The  maximum  camber  is  5/40  =  0.125. 
In  other  words,  the  maximum  depth  of  this  wing  is  12.5 
per  cent  of  the  chord,  and  unless  otherwise  specified,  is 
taken  as  being  the  camber  of  the  top  surface. 

The  maximum  camber  of  a  modern  wing  is  generally 
in  the  neighborhood  of  0.08,  although  there  are  several 
successful  sections  that  are  well  below  this  figure.  Unfor- 
tunately, the  camber  is  not  a  direct  index  to  the  value  of 
a  wing,  either  in  regard  to  lifting  ability  or  efficiency. 
By  knowing  the  camber  of  a  wing  we  cannot  directly 


132  WING  SECTIONS 

calculate  the  lift  or  drag,  for  there  are  several  examples 
of  wings  having  w^idely  different  cambers  that  give  prac- 
tically the  same  lift  and  drift.  At  the  present  time,  we 
can  only  determine  the  characteristics  of  a  wing  by  expe- 
riment, either  on  a  full  size  wing  or  on  a  scale  model. 

In  the  best  wing  sections,  the  greatest  thickness  and 
camber  occurs  at  a  point  about  0.3  of  the  chord  from  the 
front  edge,  this  edge  being  much  more  blunt  and  abrupt 
than  the  portions  near  the  trailing  edge.  An  efficient  wing 
tapers  very  gradually  from  the  point  of  maximum  camber 
towards  the  rear.  This  is  usually  a  source  of  difficulty 
from  a  structural  standpoint  since  it  is  difficult  to  get  an 
efficient  depth  of  wing  beam  at  a  point  near  the  trailing 
edge.  A  number  of  experiments  performed  by  the  Na- 
tional Physical  Laboratory  show  that  the  position  of 
maximum  ordinate  or  camber  should  be  located  at  33.2 
per  cent  from  the  leading  edge.  This  location  gives  the 
greatest  lift  per  square  foot,  and  also  the  least  resistance 
for  the  weight  lifted.  Placing  the  maximum  ordinate 
further  forward  is  worse  than  placing  it  to  the  rear. 

Thickening  the  entering  edge  causes  a  proportionate 
loss  in  efficiency.  Thickening  the  rear  edge  also  decreases 
the  efficiency  but  does  not  afifect  the  weight  lifting  value 
to  any  great  extent.  The  camber  of  the  under  surface 
seems  to  have  but  little  effect  on  the  efficiency,  but  the 
lift  increases  slightly  with  an  increase  in  the  camber  of  the 
lower  surface.  Increasing  the  camber  of  the  lower  surface 
decreases  the  thickness  of  the  wing  and  hence  decreases 
the  strength  of  the  supporting  members,  particularly  at 
points  near  the  trailing  edge.  The  increase  of  lift  due  to 
increasing  the  under  camber  is  so  slight  as  to  be  hardly 
worth  the  sacrifice  in  strength.  Variations  in  the  camber 
of  the  upper  surface  are  of  much  greater  importance.  It 
is  on  this  surface  that  the  greater  part  of  the  lift  takes 
places,  hence  a  change  in  the  depth  of  this  curve,  or  in  its 
outline,  will  cause  wider  variations  in  the  characteristics 


WIXG  SECTIONS 


133 


of  the  wing  than  would  be  the  case  with  the  under  sur- 
face. Increasing  the  upper  camber  by  about  60  per  cent 
may  double  the  lift  of  the  upper  surface,  but  the  relation 
of  the  lift  to  the  drag  is  increased.  From  this,  it  will  be 
seen  that  direct  calculations  from  the  outline  would  be 
most  difficult,  and  in  fact  a  practical  impossibility  at  the 
present  time. 

By  putting  a  reverse  curve  in  the  trailing  edge  of  a  wing, 
as  shown  by  Fig.  2,  the  stability  of  the  wing  may  be 


PR/iC  TICAL  L IMI T  QF  REFLjEX 


.  £'XC£S3IVE:  RERLEiX 

(C.  R  MO  VEMErrr  fr/^ctic/^l  l  y^£-Ro) 


Fig.  2.  (Upper)  Shows  a  Slight  "Reflex"  or  Upward  Turn  of  the  Trailing 
Edge.  Fig.  3.  (Lower)  Shows  an  Excessive  Reflex  Which  Greatly 
Reduces  the  C.P.  Movement. 


increased  to  a  surprising  degree,  but  the  lift  and  efficiency 
are  correspondingly  reduced  with  each  increase  in  the 
amount  of  reverse  curvature.  In  this  way,  stability  is 
attained  at  the  expense  of  efficiency  and  lifting  power. 
With  the  rear  edge  raised  about  0.037  of  the  chord,  the 
N.  P.  L.  found  that  the  center  of  pressure  could  be  held 
stationary,  but  the  loss  of  lift  was  about  25  per  cent  and 
the  loss  of  efficiency  amounted  practically  12  per  cent. 
With  very  slight  reverse  curvatures  it  has  been  possible 
to  maintain  the  lift  and  efficiency,  and  at  the  same  time  to 
keep  the  center  of  pressure  movement  down  to  a  reason- 
able extent.     The  New  U.  S.  A.  sections  and  the  Eiffel 


134  WING  SECTIONS 

No.  32  section  are  examples  of  excellent  sections  in  which 
a  slight  reverse  or  ''reflex"  curvature  is  used.  The  Eiffel 
32  wing  is  efficient,  and  at  the  same  the  center  of  pressure 
movement  between  incident  angles  of  0°  and  10°  is  prac- 
tically negligible.  This  wing  is  thin  in  the  neighborhood 
of  the  trailing  edge,  and  it  is  very  difficult  to  obtain  a 
strong  rear  spar. 

Wing  Selection.  No  single  wing  section  is  adapted  to 
all  purposes.  Some  wings  give  a  great  lift  but  are  ineffi- 
cient at  small  angles  and  with  light  loading.  There  are 
others  that  give  a  low  lift  but  are  very  efficient  at  the 
small  angles  used  on  'high  speed  machines.  As  before 
explained,  there  are  very  stable  sections  that  give  but  poor 
results  when  considered  from  the  standpoint  of  lift  and 
efficiency.  The  selection  of  any  one  wing  section  depends 
upon  the  type  of  machine  upon  which  it  is  to  be  used, 
whether  it  is  to  be  a  small  speed  machine  or  a  heavy 
flying  boat  or  bombing  plane. 

There  are  a  multitude  of  wing  sections,  each  possessing 
certain  admirable  features  and  also  certain  faults.  To  list 
all  of  the  wings  that  have  been  tried  or  proposed  would 
require  a  book  many  times  the  size  of  this,  and  for  this 
reason  I  have  kept  the  list  of  wings  confined  to  those  that 
have  been  most  commonly  employed  on  prominent  ma- 
chines, or  that  have  shown  evidence  of  highly  desirable 
and  special  qualities.  This  selection  has  been  made  with 
a  view  of  including  wings  of  widely  varying  character- 
istics so  that  the  data  can  be  applied  to  a  wide  range  of 
aeroplane  types.  Wings  suitable  for  both  speed  and 
weight  carrying  machines  have  been  included. 

The  wings  described  are  the  U.  S.  A.  Sections  No.  1, 
2,  3,  4,  5  and  6;  the  R.  A.  F.  Sections  Nos.  3  and  6,  and  the 
well  known  Eiffel  Wings  No.  32,  36  and  37.  The  data 
given  for  these  wings  is  obtained  from  wind  tunnel  tests 
made  at  the  Massachusetts  Institute  of  Technology,  the 
National  Physical  Laboratory  (England),  and  the  Eiffel 


WING  SECTIONS  135 

Laboratory  in  Paris.  For  each  of  these  sections  the  lift 
co-efficient  (Ky),  the  Hft-drift  ratio  (L/D),  and  the  drag 
co-efficient  (K;^)  are  given  in  terms  of  miles  per  hour  and 
pounds  per  square  foot.  Since  these  are  the  results 
for  model  wings,  there  are  certain  corrections  to  be 
made  when  the  full  size  wing  is  considered,  these  cor- 
rections being  made  necessary  by  the  fact  that  the  drag 
does  not  vary  at  the  same  rate  as  the  lift.  This  *'Size"  or 
"Scale"  correction  is  a  function  of  the  product  of  the  wing 
span  in  feet  by  the  velocity  of  the  wind  in  feet  per  second. 
A  large  value  of  the  product  results  in  a  better  wing  per- 
formance, or  in  other  words,  the  large  wing  will  always 
give  better  lift-drag  ratios  than  would  be  indicated  by  the 
model  tests.  The  lift  co-efficient  Ky  is  practically  unaf- 
fected by  variations  in  the  product.  If  the  model  tests 
are  taken  without  correction,  the  designer  will  always 
be  on  the  safe  side  in  calculating  the  power.  The  method 
of  making  the  scale  corrections  will  be  taken  up  later. 

Of  all  the  sections  described,  the  R.  A.  F.-6  is  probably 
the  best  known.  The  data  on  this  w^ing  is  most  complete, 
and  in  reality  it  is  a  sort  of  standard  by  which  the  per- 
formance of  other  wings  is  compared.  Data  has  been 
published  which  describes  the  performance  of  the 
R.  A.  F.-6  used  in  monoplane,  biplane  and  triplane  form ; 
and  with  almost  every  conceivable  degree  of  stagger, 
sweep  back,  and  decalage.  In  addition  to  the  laboratory 
data,  the  wing  has  also  been  used  with  great  success  on 
full  size  machines,  principally  of  the  ''Primary  trainer" 
class  where  an  ''All  around"  class  of  wing  is  particularly 
desirable.  It  is  excellent  from  a  structural  standpoint 
since  the  section  is  comparatively  deep  in  the  vicinity  of 
the  trailing  edge.  The  U.  S.  A.  sections  are  of  com- 
paratively recent  development  and  are  decided  improve- 
ments on  the  R.  A.  F.  and  Eiffel  sections.  The  only 
objection  is  the  limited  amount  of  data  that  is  available 
on  these  wings— limited  at  least  when  the  R.  A.  F.  data 


136  WING  SECTIONS 

is  considered — as  we  have  only  the  figures  for  the  mono- 
plane arrangement. 

WING  SELECTION. 

(1)  Lift-Drag  Ratio.  The  lift-drag  ratio  (L/D)  of  a 
wing  is  the  measure  of  wing  efficiency.  Numerically, 
this  is  equal  to  the  lift  divided  by  the  horizontal  drag, 
both  quantities  being  expressed  in  pounds.  The  greater 
the  weight  supported  by  a  given  horizontal  drag,  the  less 
will  be  the  power  required  for  the  propulsion  of  the  aero- 
plane, hence  a  high  value  of  L/D  indicates  a  desirable 
w^ing  section — at  least  from  a  power  standpoint.  In  the 
expression  L/D,  L  =  lift  in  pounds,  and  D  =  horizontal 
drag  in  pounds.  Unfortunately,  this  is  not  the  only  im- 
portant factor,  since  a  wing  having  a  great  lift-drag  is 
usually  deficient  in  lift  or  is  sometimes  structurally  weak. 

The  lift-drag  ratio  varies  with  the  angle  of  incidence 
(i),  reaching  a  maximum  at  an  angle  of  about  4°  in  the 
majority  of  wings.  The  angle  of  incidence  at  which  the 
lift-drag  is  a  maximum  is  generally  taken  as  the  angle  of 
incidence  for  normal  horizontal  flight.  At  angles  either 
greater  or  less,  the  L/D  falls  off,  generally  3.t  a  very  rapid 
rate,  and  the  power  increases  correspondingly.  Very  effi- 
cient wings  may  have  a  ratio  higher  than  L/D  =  20  at  an 
angle  of  about  4°,  while  at  16°  incidence  the  value  may 
be  reduced  to  L/D  =  4,  or  even  less.  The  lift  is  generally 
greatest  at  about  16°.  The  amount  of  variation  in  the 
lift,  and  lift-drag,  corresponding  to  changes  in  the  inci- 
dence differs  among  the  different  types  of  wings  and 
must  be  determined  by  actual  test. 

After  finding  a  wing  with  a  good  value  of  L/D,  the 
value  of  the  lift  co-efficient  Ky  should  be  determined  at 
the  angle  of  the  maximum  L/D.  With  two  wings  having 
the  same  lift-drag  ratio,  the  wing  having  the  greatest  lift 
(Ky)  at  this  point  is  the  most  desirable  wing  as  the  greater 


WING  SECTIONS  137 

lift  will  require  less  area  and  will  therefore  result  in  less 
head  resistance  and  less  weight.  Any  increase  in  the  area 
not  only  increases  the  weight  of  the  wing  surface  proper, 
but  also  increases  the  wiring  and  weights  of  the  structural 
members.  With  heavy  machines,  such  as  seaplanes  or 
bomb  droppers,  a  high  value  of  Ky  is  necessary  if  the 
area  is  to  be  kept  within  practical  limits.  A  small  fast 
scouting  plane  requires  the  best  possible  lift-drag  ratio 
at  small  angles,  but  requires  only  a  small  lift  co-efficient. 
At  speeds  of  over  100  miles  per  hour  a  small  increase  in 
the  resistance  will  cause  a  great  increase  in  the  power. 

(2)  Maximum  Lift  (Ky).  With  a  given  wing  area  and 
weight,  the  maximum  value  of  the  lift  co-efficient  (Ky) 
determines  the  slow  speed,  or  landing  speed,  of  the  aero- 
plane. The  greater  the  value  of  Ky,  the  slower  can  be  the 
landing  speed.  For  safety,  the  landing  speed  should  be 
as  low  as  possible. 

In  the  majority  of  wings,  the  maximum  lift  occurs  at 
about  16°  of  incidence,  and  in  several  sections  this  max- 
imum is  fairly  well  sustained  over  a  considerable  range 
of  angle.  The  angle  of  maximum  lift  is  variously  known 
as  the  "Stalling  angle"  or  the  "Burble  point,"  since  a 
change  of  angle  in  either  direction  reduces  the  lift  and 
tends  to  stall  the  aeroplane.  For  safety,  the  angle  range 
for  maximum  lift  should  be  as  great  as  possible,  for  if  the 
lift  falls  oft"  very  rapidly  with  an  increase  in  the  angle  of 
incidence,  the  pilot  may  easily  increase  the  angle  too  far 
and  drop  the  machine.  In  the  R.  A.  F.-3  wing,  the  lift 
is  little  altered  through  an  angle  range  of  from  14°  to 
16.5°,  the  maximum  occurring  at  15.7°,  while  with  the 
R.  A.  F.-4,  the  lift  drops  very  suddenly  on  increasing  the 
angle  above  15°.  The  range  of  the  stalling  angle  in  any 
of  the  wings  can  be  increased  by  suitable  biplane  or  tri- 
plane  arrangements.  If  large  values  of  Hft  are  accom- 
panied by  a  fairly  good  L  D  value  at  large  angles,  the 
wing  section  will  be  suitable  for  heavy  machines. 


138  WING  SECTIONS 

(3)  Center  of  Pressure  Movement.  The  center  of  pres- 
sure movement  with  varying  angles  of  incidence  is  of  the 
greatest  importance,  since  it  not  only  determines  the 
longitudinal  stability  but  also  has  an  important  effect 
upon  the  loading  of  the  wing  spars  and  ribs.  With  the 
majority  of  wings  a  decrease  in  the  angle  of  incidence 
causes  the  center  of  pressure  to  move  back  toward  the 
trailing  edge  and  hence  tends  to  cause  nose  diving.  When 
decreased  beyond  0°  the  movement  is  very  sharp  and 
quick,  the  C.  P.  moving  nearly  half  the  chord  width  in  the 
change  from  0°  to  -1.5°.  The  smaller  the  angle,  the  more 
rapid  will  be  the  movement.  Between  6°  and  16°,  the 
center  of  pressure  lies  near  a  point  0.3  of  the  chord  from 
the  entering  edge  in  the  majority  of  wing  sections.  Reduc- 
ing the  angle  from  6°  to  2°  moves  the  C.  P.  back  to 
approximately  0.4  of  the  chord  from  the  entering  edge. 

There  are  wing  sections,  however,  in  which  the  C.  P. 
movement  is  exceedingly  small,  the  Eiffel  32  being  a  not- 
able example  of  this  type.  This  wing  is  exceedingly 
stable,  as  the  C.  P.  remains  at  a  trifle  more  than  0.30  of  the 
chord  through  nearly  the  total  range  of  flight  angles.  An 
aeroplane  equipped  with  the  Eiffel  32  wing  could 
be  provided  with  exceedingly  small  tail  surfaces  with- 
out a  tendency  to  dive.  Should  the  elevator  become 
inoperative  through  accident,  the  machine  could  probably 
be  landed  without  danger.  This  wing  has  certain  objec- 
tionable features,  however,  that  offset  the  advantages. 

It  will  be  noted  that  with  the  unstable  wings  the  center 
of  pressure  movement  always  tends  to  aggravate  the 
wing  attitude.  If  the  machine  is  diving,  the  decrease  in 
angle  causes  the  C.  P.  to  move  back  and  still  further 
increase  the  diving  tendency.  If  the  angle  is  suddenly 
increased,  the  C.  P.  moves  forward  and  increases  the 
tendency  toward  stalling. 

If  the  center  of  pressure  could  be  held  stationary  at 
one  point,  the  wing  spars  could  be  arranged  so  that  each 


WING  SECTIONS  139 

Spar  would  take  its  proper  proportion  of  the  load.  As  it 
is,  either  spar  may  be  called  upon  to  carry  anywhere  from 
three-fourths  of  the  load  to  entire  load,  since  at  extreme 
angles  the  C.  P.  is  likely  to  lie  directly  on  either  of  the 
spars.  Since  the  rear  spar  is  always  shallow  and  ineffi- 
cient, this  is  most  undesirable.  This  condition  alone  to 
a  certain  extent  counterbalances  the  structural  disadvan- 
tage of  the  thin  Eififel  32  section.  Although  the  spars  in 
this  wing  must  of  necessity  be  shallow,  they  can  be 
arranged  so  that  each  spar  will  take  its  proper  share  of 
the  load  and  with  the  assurance  that  the  loading  will 
remain  constant  throughout  the  range  of  flight  angles. 
The  comparatively  deep  front  spar  could  be  moved  back 
until  it  carried  the  greater  part  of  the  load,  thus  relieving 
the  rear  spar. 

With  a  good  lift-drag  ratio,  and  a  comparatively  high 
value  of  Ky,  the  center  of  pressure  movement  should  be 
an  important  consideration  in  the  selection  of  a  wing. 
It  should  be  remembered  in  this  regard  that  the  stability 
effects  of  the  C.  P.  movement  can  be  offset  to  a  consider- 
able extent  by  suitable  biplane  arrangements. 

(4)  Structural  Considerations.  For  large,  heavy  ma- 
chines, the  structural  factor  often  ranks  in  importance 
with  the  lift-drag  ratio  and  the  lift  co-efficient.  It  is  also 
of  extreme  importance  in  speed  scouts  where  the  num- 
ber of  interplane  struts  are  to  be  at  a  minimum  and  where 
the  bending  moment  on  the  wing  spars  is  likely  to  be 
great  in  consequence.  A  deep,  thick  wing  section  per- 
mits of  deep  strong  wing  spars.  The  strength  of  a  spar 
increases  with  the  square  of  its  depth,  but  only  in  direct 
proportion  to  its  width.  Thus,  doubling  the  depth  of  the 
spar  increases  the  strength  four  times,  while  doubling  the 
width  only  doubles  the  strength.  The  increase  in  weight 
would  be  the  same  in  both  cases. 

While  very  deep  wings  are  not  usually  efficient,  when 
considered  from  the  wing  section  tests  alone,  the  total 


140  WING  SECTIONS 

efficiency  of  the  wing  construction  when  mounted  on  the 
machine  is  greater  than  would  be  supposed.  This  is  due 
to  the  lightness  of  the  spars  and  to  the  reduction  in  head 
resistance  made  possible  by  a  greater  spacing  of  the  inter- 
plane  struts.  Thus,  the  deep  wing  alone  may  have  a  low 
L/D  in  a  model  test,  but  its  structural  advantages  give 
a  high  total  efficiency  for  the  machine  assembled. 

Summary.     It  will  be  seen  from  the  foregoing  matter 
that  the  selection  of  a  wing  consists  in  making  a  series  of 
compromises   and   that   no   single   wing  section  can   be 
expected  to  fulfill  all  conditions.    With  the  purpose  of  the 
proposed  aeroplane  thoroughly  in  mind,  the  various  sec- 
tions are  taken  up  one  by  one,  until  a  wing  is  found  that 
most  usefully  compromises  with  all  of  the  conditions. 
Reducing  this  investigation  to  its  simplest  elements  we 
must  follow  the  routine  as  described  above :   (1)  Lift-drift 
ratio  and  value  of  Ky  at  this  ratio.    (2)  Maximum  value  of 
Ky  and  L/D  at  this  lift.    (3)  Center  of  pressure  movement. 
(4)  Depth  of  wing  and  structural  characteristics. 
yy^^alculations  for  Lift  and  Area.    Although  the  principles 
//  of  surface  calculations  were  described  in  the  chapter  on 
11   elementary    aerodynamics,    it    will    probably    simplify 
\\  matters  to  review  these  calculations  at  this  point.     The 
lift  of  a  wing  varies  with  the  product  of  the  area,  and  the 
velocity    squared,   this    result   being   multiplied    by    the 
co-efficient  of  lift  (Ky).     The  co-efficient  varies  with  the 
wing  section,  and  with  the  angle  of  incidence.    Stated  as 
a  formula:  L  =  KyAV^  where  A  =  area  in  square  feet, 
and  V  =  velocity  of  the  wing  in  miles  per  hour.    Assum- 
ing an  area  of  200  square  feet,  a  velocity  of  80  miles  per 
hour,  and  with  K  =  0.0025,  the  total  life  (L)  becomes^ 
L=  Ky  A  V^  =0.0025  X  200  X  (80  X  80  =  3,200  pounds. 
Assuming  a  lift-drag  ratio  of  16,  the  "drag"  of  the  wing, 
or  its  resistance  to  horizontal  motion,  will  be  expressed 
by  D  =  L/r  =  3,200/16  =  200  pounds,  where  r  =  lift-drag 
ratio.    It  is  this  resistance  of  200  pounds  that  the  motor 


WING  SECTIONS  141 

must  overcome  in  driving  the  wings  through  the  air.  The 
total  resistance  offered  by  the  aeroplane  will  be  equal  to 
the  sum  of  the  wing  resistance  and  the  head  resistance  of 
the  body,  struts,  wiring  and  other  structural  parts.  In  the 
present  instance  we  will  consider  only  the  resistance  of 
the  wings. 

When  the  lift  co-efficient,  speed,  and  total  lift  are  known, 
the  area  can  be  found  from  A  =  L/KyV-,  the  lift,  of 
course,  being  taken  as  the  total  weight  of  the  machine. 
The  area  of  the  supporting  surface  for  a  speed  of  60  miles 
per  hour,  total  weight  of  2,400  pounds,  and  a  lift  co-effi- 
cient of  0.002  is  calculated  as  follows : 

A  ==  L/K,\-^=  2,400/0.002  X  (60  X  60)  =333  sq.  ft. 

A  third  variation  in  the  formula  is  that  used  in  finding 
the  value  of  the  lift  co-efficient  for  a  particular  wing  load- 
ing. From  the  weight,  speed  and  area,  we  can  find  the 
co-efficient  Ky,  and  with  this  value  we  can  find  a  wing 
that  will  correspond  to  the  required  co-efficient.  This 
method  is  particularly  convenient  when  searching  for  the 
section  with  the  greatest  lift-drag  ratio.  Ky  =  L/  AV", 
or  when  the  loading  per  square  foot  is  known,  the  co-effi- 
cient becomes  Ky=LyV".  For  example,  let  us  find  the 
co-efficient  for  a  wing  loading  of  5  pounds  per  square  foot 
at  a  velocity  of  80  miles  per  hour.  Inserting  the  numerical 
values  into  the  equation  we  have,  Ky  =  L7V"=  5/(80  X 
80)  =0.00078.  Any  wing,  at  any  angle  that  has  a  lift 
co-efficient  equal  to  0.00078  will  support  the  load  at  the 
given  speed,  although  many  of  the  wings  would  not  give 
a  satisfactory  lift-drag  ratio  with  this  co-efficient. 

It  should  be  noted  in  the  above  calculations  that  no 
correction  has  been  made  for  ''Scale,"  aspect  ratio  or 
biplane  interference.  In  other  words,  we  have  assumed 
the  figures  as  applying  to  model  monoplanes.  In  the 
following  tables  the  lift,  lift-drag  and  drag  must  be  cor- 
rected, since  this  data  was  obtained  from  model  tests  on 
monoplane  sections.     The  effects  of  biplane  interference 


142  WING  SECTIONS 

will  be  described  in  the  chapter  on  "Biplane  and  Tri- 
plane  Arrangement,"  but  it  may  be  stated  that  superposing 
the  planes  reduces  both  the  lift  co-efficient  and  the  lift- 
drag  ratio,  the  amount  of  reduction  depending  upon  the 
relative  gap  between  the  surfaces.  Thus  with  a  gap  equal 
to  the  chord,  the  lift  of  the  biplane  surface  will  only  be 
about  80  per  cent  of  the  lift  of  a  monoplane  surface  of 
the  same  area  and  section. 

Wing  Test  Data.  The  data  given  in  this  chapter  is  the 
result  of  wind  tunnel  tests  made  under  standard  condi- 
tions, the  greater  part  of  the  results  being  published  by 
the  Massachusetts  Institute  of  Technology.  The  tests 
were  all  made  on  the  same  size  of  model  and  at  the  same 
wind  speed  so  that  an  accurate  comparison  can  be  made 
between  the  different  sections.  All  values  are  for  mono- 
plane wings  with  an  aspect  ratio  of  6,  the  laboratory 
models  being  18x3  inches.  The  exception  to  the  above 
test  conditions  will  be  found  in  the  tables  of  the  Eiffel 
37  and  36  sections,  these  figures  being  taken  from  the 
results  of  Eiffel's  laboratory.  The  Eiffel  models  were 
35.4x5.9  inches  and  were  tested  at  wind  velocities  of  22.4, 
44.8,  and  67.2  miles  per  hour.  The  tests  made  at  M.  I.  T. 
were  all  made  at  a  wind  speed  of  30  miles  per  hour.  The 
lift  co-efficient  Ky  is  practically  independent  of  the  wing 
size  and  wind  velocity,  but  the  drag  co-efficient  Kx  varies 
with  both  the  size  and  wind  velocity,  and  the  variation 
is  not  the  same  for  the  different  wings.  The  results  of 
the  M.  I.  T.  tests  were  published  in  ^'Aviation  and  Aero- 
nautical Engineering"  by  Alexander  Klemin  and  G.  M. 
Denkinger. 

The  R.  A.  F.  Wing  Sections.  These  wings  are  prob- 
ably the  best  known  of  all  wings,  although  they  are 
inferior  to  the  new  U.  S.  A.  sections.  They  are  of  English 
origin,  being  developed  by  the  Royal  Aircraft  Factory 
(R.  A.  F.),  with  the  tests  performed  by  the  National 
Physical     Laboratory    at    Teddington,    England.      The 


WIXG  SECTIONS  U3 

R.  A.  F.-6  is  the  nearest  approach  to  the  all  around  wing, 
this  section  having  a  fairly  high  L/D  ratio  and  a  good 
value  of  Ky  for  nearly  all  angles.  It  is  by  no  means  a 
speed  wing  nor  is  it  suitable  for  heavy  machines,  but  it 
comprises  well  between  these  limits  and  has  been  exten- 
sively used  on  medium  size  machines,  such  as  the  Cur- 
tiss  JN4-B,  the  London  and  Provincial,  and  others.  The 
R.  A.  F.-3  has  a  very  high  value  for  Ky,  and  a  very  good 
lift-drag  ratio  for  the  high-lift  values.  It  is  suitable  for 
seaplanes,  bomb  droppers  and  other  heavy  machines  of  a 
like  nature  that  fly  at  low  or  moderate  speeds. 

The  outlines  of  these  wings  are  shown  by  Figs.  7  and  8, 
and  the  camber  ordinates  are  marked  as  percentages  of 
the  chord.  In  laying  out  a  wing  rib  from  these  diagrams, 
the  ordinate  at  any  point  is  obtained  by  multiplying  the 
chord  length  in  inches  by  the  ordinate  factor  at  that 
point.  Referring  to  the  R.  A.  F.-3  diagram,  Fig.  8,  it  will 
be  seen  that  the  ordinate  for  the  upper  surface  at  the 
third  station  from  the  entering  edge  is  0.064.  If  the  chord 
of  the  wing  is  60  inches,  the  height  of  the  upper  curve 
measured  above  the  datum  line  X-X  at  the  third  station 
will  be,  0.064X60^=3.84  inches.  At  the  same  station, 
the  height  of  the  lower  curve  will  be,  0.016  X  60  =  0.96 
inch. 

The  chord  is  divided  into  10  equal  parts,  and  at  the 
entering  edge  one  of  the  ten  parts  is  subdivided  so  as  to 
obtain  a  more  accurate  curve  at  this  point.  In  some  wing 
sections  it  is  absolutely  necessary  to  subdivide  the  first 
chord  division  as  the  curve  changes  very  rapidly  in  a 
short  distance.  The  upper  curve,  especially  at  the  enter- 
ing edge,  is  by  far  the  most  active  part  of  the  section  and 
for  this  reason  particular  care  should  be  exercised  in 
getting  the  correct  outline  at  this  point. 

Aerodynamic  Properties  of  the  R.  A.  F.  Sections.  Table 
1  gives  the  values  of  Ky,  K^,  L/D,  and  the  center  of  pres- 
sure movement  (C.  P.)  for  the  R.  A.  F.-3  section  through 


144 


WING  SECTIONS 


a  range  of  angles  varying  from  — 2°  to  20°.  The  first 
column  at  the  left  gives  the  angles  of  incidence  (i),  the 
corresponding  values  for  the  lift  (Ky)  and  the  drag  (Kx) 
being  given  in  the  second  and  third  columns,  respectively. 
The  fourth  column  gives  the  lift-drag  ratio  (L/D).  The 
fifth  and  last  column  gives  the  location  of  the  center  of 
pressure  for  each  different  angle  of  incidence,  the  figure 


(7)BJir-6    Wtn^ 


(<5)  -RAV-3  Wm& 


Figs.   7-8.     R.A.F.  Wing  Sections.     Ordinates  as  Decimals  of  the  Chord. 


indicating  the  distance  of  the  C.  P.  from  the  entering  edge 
expressed  as  a  decimal  part  of  the  chord. 

As  an  example  in  the  use  of  the  table,  let  it  be  required 
to  find  the  Hft  and  drag  of  the  R.  A.  F.-3  section  when 
inclined  at  an  angle  of  6°  and  propelled  at  a  speed  of  90 
miles  per  hour.    The  assumed  area  will  be  300  square  feet. 

At  6°  it  will  be  found  that  the  lift  co-efficient  Ky  is 
0.002369.  From  our  formulae,  the  lift  will  be  :  L  =  Ky  AV- 
er numerically,  L  =  0.002369  X  300  X  (90  X  90  =  5,756.7 
lbs.    At  the  same  speed,  but  with  the  angle  of  incidence 


WING  SECTIONS  1-15 

reduced  to  2^  the  lift  will  be  reduced  to  L  =  0.001554  X 
300  X  (90  X  90)  =3,776.2  pounds,  where  0.001554  is  the 
lift  co-efficient  at  2°.  It  will  be  noted  that  the  maximum 
lift  co-efficient  occurs  at  14°  and  continues  at  this  value 
to  a  little  past  15°.  The  lift  at  the  stalling  angle  is  fairly 
constant.from  12°  to  16°. 

The  value  of  the  drag  can  be  found  in  either  of  two 
ways:  (1)  by  dividing  the  total  lift  (L)  by  the  lift-drag 
ratio,  or  (2)  by  figuring  its  value  by  the  formula  D  = 
KxAV".  The  first  method  is  shorter  and  preferable.  By 
consulting  the  table,  it  will  be  seen  that  the  L/D  ratio  at 
6°  is  14.9.  The  total  wing  drag  will  then  be  equal  to 
5,756.7/14.9=386.4  lbs.  Figured  by  the  second  method, 
the  value  of  K^  at  6°  is  0.000159,  and  the  drag  is  therefore : 
D==K,AV2  =  0.000 159  X  300  X  (90  X  90)  =386.4.  This 
checks  exactly  with  the  first  method.  The  lift-drag  ratio 
is  best  at  4°,  the  figure  being  15.6,  while  the  lift  at  this 
point  is  0.001963.  With  the  same  area  and  speed,  the  total 
lift  of  the  surface  at  the  angle  of  best  lift-drift  ratio  will 
be  0.001963  X  300  X  (90  X  90)  =  4,770  lbs. 

At  4°  the  center  of  pressure  is  0.385  of  the  chord  from 

Table  1. 
R.A.F.  3  Wing. 

i  Ky  Kx  L/D  C.P. 

_2  000601  .000125 

—1  000879  .000109 

0  001100  .000101 

-fl  001314  .000100 

2  001554  .000105 

4  001963  .000126 

6  002369  .000159 

8  002777  .000207 

12  003439  .000315 

14  003481  .000378 

14K- 003481  .000405 

15  003481  .000425 

16  003472  .000465 

18  003406  .000598 

20  003376  .000908 


4.8 

.785 

8.0 

.620 

11.0 

.522 

13.2 

.470 

14.9 

.435 

15.6 

.385 

14.9 

.352 

13.4 

.332 

10.9 

.315 

9.2 

.298 

8.6 

.293 

8.2 

.290 

7.4 

.295 

5.7 

.328 

3.7 

.382 

146  WING  SECTIONS 

Table  2. 

R.A.F.  6. 

i                                               Ky                    Kx  L/D  C.P. 

—2  000172  .000090  0.92 

—1  000285  .000081            3.40  .682 

0  000571  .000077            7.20  .522 

1  000821  .000069  12.10  .445 

2  001072  .000068  15.75  .401 

4  001477  .000090  16.58  .360 

6  001873  .000128  14.14  .328 

8  002268  .000167  13.69  .310 

10     002634  .000207  12.92  .298 

12     002882  .000255  11.18  .289 

14     003018  .000321  9.28  .292 

16     003045  .000434  6.92  .300 

18     002987  .000698  4.48  .319 

20     002871  .000887  3.20  .360 

the  entering  edge.  If  the  chord  is  60  inches  wide,  the 
center  of  pressure  will  be  located  at  0.385  X  60  =  23.1 
inches  from  the  entering  edge.  At  15°,  the  center  of  pres- 
sure will  be  0.29X60  =  17.4  inches  from  the  entering 
edge,  or  during  the  change  from  4°  to  15°  the  center  of 
pressure  will  have  moved  forward  by  5.7  inches.  At  — 2°, 
the  pressure  has  moved  over  three-quarters  of  the  way 
toward  the  trailing  edge  — 0.785  of  the  chord,  to  be  exact 
Through  the  ordinary  flight  angles  of  from  2°  to  12°,  the 
travel  of  the  center  of  pressure  is  not  excessive. 

The  maximum  lift  co-efficient  (Ky)  is  very  high  in  the 
R.  A.  F.-3  section,  reaching  a  maximum  of  0.003481  at  an 
incidence  of  14°.  This  is  second  to  only  one  other  wing, 
the  section  U.  S.  A.-4.  This  makes  it  suitable  for  heavy 
seaplanes. 

Table  2  gives  the  aerodynamic  properties  of  the  R.  A. 
F.-6  wing,  the  table  being  arranged  in  a  manner  similar 
to  that  of  the  R.  A.  F.-3.  In  glancing  down  the  column  of 
lift  co-efficients  (Ky),  and  comparing  the  values  with 
those  of  the  R.  A.  F.-3  section,  it  will  be  noted  that  the 
lift  of  R.  A.  F.-6  is  much  lower  at  every  angle  of  incidence, 
but  that  the  lift-drag  ratio  of  the  latter  section  is  not 
always  correspondingly  higher.     At  every  angle  below 


WING  SECTIONS  147 

2°,  at  6°,  and  at  angles  above  14°,  the  L/D  ratio  of  the 
R.  A.  F.-3  is  superior  in  spite  of  its  greater  lift.  The 
maximum  L  D  ratio  of  the  R.  A.  F.-6  at  4°  is  16.58,  which 
is  considerably  higher  than  the  best  L/D  ratio  of  the 
R.  A.  F.-3.  The  best  lift  co-efficient  of  the  R.  A.  F.-6. 
0.003045,  is  very  much  lower  than  the  maximum  Ky  of 
the  R.  A.  F.-3. 

The  fact  that  the  L  D  ratio  of  the  R.  A.  F.-3  wing  is 
much  greater  at  high  lift  co-efficients,  and  large  angles 
of  incidence,  makes  it  very  valuable  as  at  this  point  the 
greater  L/D  does  not  tend  to  stall  the  plane  at  slow 
speed.  A  large  L/D  at  -great  angles,  together  with  a 
Avide  stalling  angle  tends  for  safety  in  slow  speed  flying. 
Both  wing  sections  are  structurally  excellent,  being 
very  deep  in  the  region  of  the  rear  edge,  the  R.  A.  F.-6 
being  particularly  deep  at  this  point.  A  good  deep  spar 
can  be  placed  at  almost  any  desirable  point  in  the  R.  A. 
F.-6,  and  the  trailing  edge  is  deep  enough  to  insure  against 
rib  weakness  even  with  a  comparatively  great  overhang. 
Scale  corrections  for  the  full  size  R.  A.  F.  wings  are 
very  difficult  to  make.  According  to  the  N.  P.  L.  reports, 
the  corrected  value  for  the  maximum  L/D  of  the  R.  A.  F.-3 
wing  is  18.1,  the  model  test  indicating  a  maximum  value 
of  15.6.  I  believe  that  L/D  =17.5  would  be  a  safe  full 
size  value  for  this  section.  The  same  reports  give  the 
full  size  L/D  for  the  R.  A.  F.-6  as  18.5,  which  would  be 
probably  safe  at  18.0  under  the  new  conditions. 

Properties  of  the  Eiffel  Sections  (32-36-37).  Three  of 
the  Eiffel  sections  are  shown  by  Figs.  10,  11  and  12,  these 
sections  being  selected  out  of  an  enormous  number  tested 
in  the  Eiffel  laboratories.  They  dift'er  widely,  both  aero- 
dynamically  and  structurally,  from  the  R.  A.  F.  aerocurves 
just  illustrated. 

Eiffel  32  is  a  very  stable  wing,  as  has  already  been 
pointed  out,  but  the  value  of  the  maximum  L/D  ratio 
is  in  doubt  as  this  quantity  is  very  susceptible  to  changes 


148 


WING  SECTIONS 


in  the  wind  velocity — much  more  than  in  the  average 
wing.  Since  Eiffel's  tests  were  carried  out  at  much  higher 
velocity  than  at  the  M.  I.  T.,  his  lift-drift  values  at  the 
higher  speeds  were  naturally  much  better  than  those 
obtained  by  the  American  Laboratory.  When  tested  at 
67.2  miles  per  hour  the  lift-drift  ratio  for  the  Eiffel  32 


Eirr£JljI\fo.32. 


BrrFEJL  No.37 


0.067^ 


Figs.   10-11-12.     Ordinates  for  Three  Eiffel  Wing  Sections. 


was  18.4  while  at  22.4  miles  per  hour,  the  ratio  dropped 
to  13.4.  This  test  alone  will  give  an  idea  as  to  the  varia- 
tion possible  with  changes  in  scale  and  wind  velocity.  The 
following  table  gives  the  results  of  tests  carried  out  at 
the  Massachusetts  laboratory,  reported  by  Alexander 
Klemin  and  G.  M.  Denkinger  in  "Aviation  and  Aero- 
nautical Engineering."     Wind  speed,  30  miles  per  hour. 


WING  SECTIONS  U9 

Table  3. 

Eiffel  32. 

i                                             Ky                   Kx  L/D  C.R 

—2 +.000159  .0000735  2.2  .330 

— 1 000378  .0000765  4.9  .327 

0 000591  .0000747  7.9  .320 

1 000798  .0000785  10.2  .318 

2 001011  .0000843  12.0  .310 

4 001467  .0001059  13.9  .305 

6 001894  .0001305  14.5  .304 

8 002250  .0001741  12.9  .308 

12 002732  .000324  8.4  .335 

16 002908  .000710  4.1  .357 

18 002761  .000846  3.3  .370 

20 002642  .000956  2.8  .378 

The  C.  P.  Travel  in  the  Eiffel  wing  is  very  small, 
as  will  be  seen  from  Table  3.  At  —2°  the  C.  P.  is  0.33 
of  the  chord  from  the  leading  edge  and  only  moves  back 
to  0.378  at  an  angle  of  20^,  the  intermediate  changes 
being  very  gradual,  reaching  a  minimum  of  0.304  at  6°  in- 
cidence. The  maximum  Ky  of  Eififel  32  is  0.002908,  while 
for  the  R.  A.  F.-6  wing,  Ky^  0.003045  maximum,  both 
co-efficients  being  a  maximum  at  16°  incidence,  but  the 
lift-drag  at  maximum  Ky  is  much  better  for  the  R.  A.  F.-6. 

Structurally,  the  EifTel  32  is  at  a  disadvantage  when 
compared  with  the  R.  A.  F.  sections  since  it  is  very  nar- 
row at  points  near  the  trailing  edge.  This  would  neces- 
sitate moving  the  rear  spar  well  up  toward  the  center  with 
the  front  spar  located  very  near  the  leading  edge.  This 
is  the  type  of  wing  used  in  a  large  number  of  German 
machines.  It  will  also  be  noted  that  there  is  a  very  pro- 
nounced reverse  curve  or  "Reflex"  in  the  rear  portion, 
the  trailing  edge  actually  curving  up  from  the  chord  line. 

Eiffel  36  is  a  much  thicker  wing  than  either  of  the  other 
Eiffel  curves  shown,  and  is  deficient  in  most  aerodynam- 
ical respects.  It  has  a  low  value  for  Ky  and  a  poor  lift- 
drag  ratio.  It  has,  however,  been  used  on  several  Amer- 
ican training  machines,  probably  for  the  reason  that  it 
permits  of  sturdy  construction. 


150 


WING  SECTIONS 


Eiffel  Z7  is  essentially  a  high-speed  wing  having  a  high 
L/D  ratio  and  a  small  lift  co-efficient.  The  maximum 
lift-drag  ratio  of  20.4  is  attained  at  a  negative  angle  — 0.8°. 
The  value  of  Ky  at  this  point  is  0.00086,  an  extremely  low 


34.0 


4) 


»li 


.00300 
290 

.00230 
210 
260  -.00260 
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-^"     e"    a"    /<9°    z^"   z^"   /e' 

ANGLE  OF  INC/OENCE 


Fig.    13.      Characteristic   Curves   for   Eiffel   Wing   Sections. 


figure.  The  maximum  Ky  is  0.00288  at  14.0°,  the  L/D 
ratio  being  4.0  at  this  angle.  Structurally  it  is  the  worst 
wing  that  we  have  yet  discussed,  being  almost  "paper 
thin"  for  a  considerable  distance  near  the  trailing  edge. 
The  under  surface  is  deeply  cambered,  with  the  maximum 
under  camber  about  one-third  from  the  trailing  edge.    It 


WING  SECTIONS  151 

is  impossible  to  use  this  wing  without  a  very  long  over- 
hang in  the  rear  of  the  section,  and  like  the  Eiffel  32,  the 
front  spar  must  be  very  far  forward.  For  those  desiring 
flexible  trailing  edges,  this  is  an  ideal  section.  This  wing 
is  best  adapted  for  speed  scouts  and  racing  machines 
because  of  its  great  L/D,  but  as  its  lift  is  small  and  the 
center  of  pressure  movement  rapid  at  the  point  of  max- 
imum lift-drag,  it  would  be  necessary  to  fly  at  a  small 
range  of  angles  and  land  at  an  extremely  high  speed. 
Any  slight  change  in  the  angle  of  incidence  causes  the 
lift-drag  ratio  to  drop  at  a  rapid  rate,  and  hence  the  wing 
could  only  be  manipulated  at  its  most  efficient  angle  by  an 
experienced  pilot.  Again,  the  angle  of  maximum  L/D  is 
only  a  few  degrees  from  the  angle  of  no  lift. 

U.  S.  A.  Wing  Sections.  These  wing  sections  were  de- 
veloped by  the  Aviation  Section  of  the  Signal  Corps, 
United  States  Army,  and  are  decided  improvements  on 
any  wing  sections  yet  published.  The  six  U.  S.  A.  wings 
cover  a  wide  range  of  application,  varying  as  they  do, 
from  the  high  speed  sections  to  the  heavy  lift  wings  used 
on  large  machines.  The  data  was  first  published  by  Cap- 
tains Edgar  S.  Gorrell  and  H.  S.  Martin,  U.  S.  A.,  by  per- 
mission of  Professor  C.  H.  Peabody,  Massachussetts  In- 
stitute of  Technology.  An  abstract  of  the  paper  by  Alex- 
ander Klemin  and  T.  H.  Huff  was  afterwards  printed  in 
''Aviation  and  Aeronautical  Engineering."  While  several 
of  the  curves  are  modifications  of  the  R.  A.  F.  sections 
already  described,  they  are  aerodynamically  and  structur- 
ally superior  to  the  originals,  and  especial  attention  is 
called  to  the  marked  structural  advantages. 

U.  S.  A.-l  and  U.  S.  A. -6  are  essentially  high  speed  sec- 
tions with  a  very  high  lift-drag  ratio,  these  wings  being 
suitable  for  speed  scouts  or  pursuit  machines.  The  dif- 
ference between  the  wings  is  very  slight,  U.  S.  A.-l  with 
Kj.^=0.00318  giving  a  better  landing  speed,  while  U.  S.  A.-6 
is  slightly  more  efficient  at  low  angles  and  high  speeds. 


152 


WING  SECTIONS 


E  ^     i     ^     ^ 


•1.00 


0.10:^0.1 

I 


■O.I(M— O.IO-^O. 

fr 
AS 


USA-l 


Fig.  14.  U.S.A.  Wing  Sections  Nos.  1-2-3-4-S-6,  Showing  the  Ordinates  at 
the  Various  Stations  Expressed  as  Decimals  of  the  Chord.  U.S.A.- 
4  Is  a  Heavy  Lift  Section,  While  U.S.A.-l  and  U.S.A.-6  are  High 
Speed  Wings.  For  Any  Particular  Duty,  the  Above  Wings  Are 
Very  Deep  and  Permit  of  Large  Structural  Members.  The  Center 
of  Pressure  Movement  Is  Comparatively  Slight. 


WING  SECTIONS  153 

With  0-  incidence,  the  ratio  of  U.  S.  A.-l=11.0  while  the 
lift-drag  of  U.  S.  A.-6  at  0°  incidence  is  13.0.  The  maxi- 
mum lift  of  U.  S.  A.-l  is  superior  to  that  of  Eiffel  32,  and 
the  maximum  lift-drag  ratio  at  equal  speeds  is  far  superior, 
being  17.8  against  14.50  of  the  Eiffel  32.  Compared  with 
the  Eiffel  32  it  will  be  seen  that  the  U.  S.  A.  sections  are 
far  better  from  a  structural  point  of  view,  especially  in  the 
case  of  U.  S.  A.-l.  The  depth  in  the  region  of  the  rear 
spar  is  exceptionally  great,  about  the  same  as  that  of  the 
R.  A.  F.-6.  \\'hile  neither  of  the  U.  S.  A.  wings  are  as 
stable  as  the  Eift'el  32,  the  motion  of  the  C.  P.  is  not  sudden 
nor  extensive  at  ordinary  flight  angles. 

Probably  one  of  the  most  remarkable  of  the  United 
States  Army  wings  is  the  U.  S.  A.-4  which  has  a  higher 
maximum  lift  co-efficient  (Ky)  than  even  the  R.  A.  F.-3. 
The  maximum  Ky  of  the  U.  S.  A.-4  is  0.00364  compared 
with  the  R.  A.  F.-3  in  which  Ky  (Maximum)=^.003481. 
Above  4°  incidence,  the  lift-drag  ratio  of  the  U. 
S.  A.-4  is  generally  better  than  that  of  the  R.  A.  F.-3,  the 
maximum  L  D  at  4°  being  considerably  better.  This  is  a 
most  excellent  wing  for  a  heavy  seaplane  or  bomber.  The 
U.  S.  A.-2  has  an  upper  surface  similar  to  that  of  the  R. 
A.  F.-3,  but  the  wing  has  been  thickened  for  structural 
reasons,  thus  causing  a  modification  in  the  lower  surface. 
This  results  in  no  particular  aerodynamic  loss  and  it  is 
much  better  at  points  near  the  rear  edge  for  the  reception 
of  a  deep  and  efficient  rear  spar. 

U.  S.  A.-3  is  a  modification  of  U.  S.  A.-2,  and  like  U.  S. 
A.-2  would  fall  under  the  head  of  "All  around  wings,"  a 
type  similar,  but  superior  to  R.  A.  F.-6.  These  wings  are 
a  compromise  between  the  high  speed  and  heavy  lift 
types — suitable  for  training  schools  or  exhibition  flyers. 
Both  have  a  fairly  good  L/D  ratio  and  a  corresponding 
value  for  Ky. 

U.  S.  A. -5  has  a  very  good  maximum  lift-drag  ratio 
(16.21)  and  a  good  lift-drag  ratio  at  the  maximum  Ky. 


154  WIXG  SECTIONS 

Its  maximum  Ky  is  superior  to  all  sections  with  the  ex- 
ception of  U.  S.  A. -2  and  4.  Structurally  it  is  very  good, 
being  deep  both  fore  and  aft. 

In  review  of  the  U.  S.  A.  sections,  it  may  be  said  that 
they  are  all  remarkable  in  having  a  very  heavy  camber  on 
both  the  upper  and  lower  surfaces,  and  at  the  same  time 
are  efficient  and  structurally  excellent.  This  rather  con- 
tradicts the  usual  belief  that  a  heavy  camber  will  produce 
a  low  lift-drag  ratio,  a  belief  that  is  also  proven  false  by 
the  excellent  performance  of  the  Eiffel  Z7  section.  The 
maximum  Ky  is  also  well  sustained  at  and  above  0.003. 
There  is  no  sharp  drop  of  lift  at  the  "Stalling  angle"  and 
the  working  range  of  incidence  is  large. 

Curtiss  Wing  and  Double  Cambered  Sections.  An  old 
type  of  Curtiss  wing  is  shown  by  Fig.  15.  It  is  very  thick 
and  an  efficient  wing  for  general  use.  It  will  be  noticed 
that  there  is  a  slight  reflex  curve  at  the  trailing  edge  of 
the  under  surface  and  that  there  is  ample  spar  room  at 
almost  any  point  along  the  section.  The  nose  is  very 
round  and  thick  for  a  wing  possessing  the  L/D  charac- 
teristics exhibited  in  the  tests.  The  conditions  of  the 
test  were  the  same  as  for  the  preceding  wing  sections. 

Fig.  16  shows  a  remarkable  Curtiss  section  designed 
for  use  as  a  stabilizing  surface.  It  is  double  cambered, 
the  top  surface  being  identical  with  the  lower,  and  is  there- 
fore non-lifting  with  the  chord  horizontal.  The  force  ex- 
erted by  the  surface  is  equal  with  equal  positive  or  nega- 
tive angles  of  incidence,  a  valuable  feature  in  a  control 
surface.  "  In  spite  of  its  great  thickness,  it  is  of  excellent 
stream  line  form  and  therefore  has  a  very  good  lift-drag 
ratio.  At  0°  angle  of  incidence  the  resistance  is  at  a  mini- 
mum, and  is  much  less  than  that  of  a  thin,  square  edged, 
flat  plate.  This  double  cambered  plane  reduces  the  stay 
bracing  and  head  resistance  necessary  with  the  flat  type 
of  stabilizer  surface. 


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WING  SECTIONS  Wl 

The  Curtiss  sections  mentioned  above  were  described 
in    "Aviation    and    Aeronautical    Engineering"    by    Dr. 


Os)  CUETJSS  Wm^  .SECTION. 


^oz^iy^ 


my/MM^.wmm. 


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Oe)  OuHTissJDouBLE  Camber. 


Fig.   15.     Old  Type  of  Curtiss  Wing.      16.   Curtiss  Double  Camber  for  Con- 
trol Surfaces. 


CURTISS  WING  SECTION. 

(i)                  Kr  Kx 

0° 0.000485  0.000060 

2° 000948  .000068 

4° 001435  .000087 

6° 001750  .000117 

8° 002040  .000168 

10° 002280  .000265 

12° 002420  .000421 

14° 002390  .000548 

16° 002300  .000688 

20° 002280  .000845 


CURTISS  DOUBLE  CAAIBERED  SECTION 

(i)                                                           Ky  Kx 

0° 0.000000  0.0000150 

2° 000170  .0000182 

3° 000250  .0000205 

4° 000318  .0000245 

6° 000455  .0000400 

8= 000592  .0000625 

10° 000725  .0001040 

12° 000845  .0001680 

14° 000882  .0002320 

16° 000885  .0002800 


L/D 

C.  P. 

8.00 

0.545 

14.42 

.468 

17.45 

.421 

15.00 

.382 

12.10 

.365 

8.60 

.350 

5.75 

.335 

4.21 

.345 

3.45 

.365 

2.70 

.425 

CTION. 
L/D 

C  P. 

00.00 

0.000 

10.65 

.210 

13.00 

.212 

12.70 

.215 

11.15 

.218 

9.50 

.222 

7.25 

.225 

5.00 

.255 

3.42 

.302 

1.11 

.370 

158  WING  SECTIONS 

Jerome  C.  Hunsaker,  but  the  figures  in  the  above  table 
were  obtained  by  the  author  on  a  sliding  test  wire  arrange- 
ment that  has  been  under  development  for  some  time. 
At  the  time  of  writing  several  of  the  U.  S.  A.  sections  are 
under  investigation  on  the  same  device. 

CORRECTION  FACTORS  FOR  WING  FORM 
AND  SIZE. 

Aspect  Ratio.  As  previously  explained,  the  aspect 
ratio  is  the  relation  of  the  span  to  the  chord,  and  this  ratio 
has  a  considerable  effect  upon  the  performance  of  a  wing. 
In  the  practical  full  size  machine  the  aspect  ratio  may 
range  from  5  in  monoplanes,  and  small  biplanes,  to  10  or 
12  in  the  larger  biplanes.  The  aspect  in  the  case  of  tri- 
planes  is  even  greater,  some  examples  of  the  latter  having 
aspects  of  16  to  20.  In  general,  the  aspect  ratio  increases 
with  the  gross  Aveight  of  the  machine.  Control  surfaces, 
such  as  the  rudder  and  elevator,  usually  have  a  much 
lower  aspect  ratio  than  the  main  lifting  surfaces,  particu- 
larly when  flat  non-lifting  control  surfaces  are  used.  The 
aspect  of  elevator  surfaces  will  range  from  unity  to  3, 
while  the  vertical  rudders  generally  have  an  aspect  of  1. 

With  a  given  w^ing  area,  the  span  increases  directly 
with  an  increase  in  the  aspect  ratio.  The  additional 
weight  of  the  structural  members  due  to  an  increased 
span  tend  to  offset  the  aerodynamic  advantages  gained  by 
a  large  aspect  ratio,  and  the  increased  resistance  due  to 
the  number  and  size  of  the  exposed  bracing  still  further 
reduces  the  advantage. 

Effects  of  Aspect  Ratio.  Variations  in  the  aspect  ratio 
do  not  give  the  same  results  in  all  wing  sections,  and  the 
lift  co-efficient  and  L/D  ratio  change  in  a  very  irregular 
manner  with  the  angle  of  incidence.  The  following 
tables  give  the  results  obtained  by  the  N.  P.  L.  on  a 
Bleriot    wing    section,    the    aspect    ratio    being    plotted 


WING  SECTIONS  159 

against  the  angle  of  incidence.  The  figures  are  com- 
parative, an  aspect  factor  of  unity  (1.000)  being  taken  for 
an  aspect  ratio  of  6  at  each  angle  of  incidence.  To  obtain 
an  approximation  for  any  other  wing  section  at  any  other 
aspect  ratio,  multiply  the  model  test  (Aspect  =  6)  by  the 
factor  that  corresponds  to  the  given  angle  and  aspect 
ratio.  At  the  extreme  right  of  the  table  is  a  column  of 
rough  averages,  taken  without  regard  to  the  angles. 

EFFECT  OF  ASPECT  RATIO   ON  COEFFICIENT  Ky 


ASPECT 

WERAGE 
VAI.rE 

RATIO 

6°        2°        4°     "6°        8°       10°      12°      14°      16° 

3 

0.961 

4 

i  293  1.009  l.OOO  6.930  6.924  6.965  6.892  6.888  6.982 

0.966 

5 

1009  0.938  0.944  0.952  0.918  0.970  0.972  0.957  0.992 

0.972 

6 

1000  1000  1.000   1.000  1.000  1.000  1.000  1.000  1.000 

1.000 

7 

1073  1.033  1.038  0.990  0.953  0.963  0.977  0.977  0.989 

0.988 

8 

1.174   1.047  1.031  1.070  1.040  1.045  1.038  1.000  0.992 

1.062 

FFFFCT  OF  ASPECT  RATIO  ON  L/D  RATIO 

ASPECT 

AVERAGE 
VALUE 

RATIO                                                                                    .,-x™     ^™     TXTr,,r.r.XTr^T- 

3 

4 

0°        2°        4°        6°        8°       10°      12°      14°      16^ 

0.722 

1285  6  896  6.826  6.829  6.8i9  6.835  6.864  6.*838  6.856 

0.820 

5 

l"058  0  953  0.913  0.933  0.900  0.912  0.944  0.928  0.989 

0.910 

6 

1  000  1  000  1.000   1.000  1.000  1.000  1.000  1.000  1.000 

1.000 

7 

1000  1047  1.060   1.112  1.063  1.105  1.084  1.040  1.024 

1.080 

8 

1173  1.036  1.029  1.157  1.149  1.156  1.140   1.101  0.989 

1.111 

The  column  of  average  values  is  not  the  average  of  the 
tabular  values  but  is  the  average  of  the  results  obtained 
by  a  number  of  investigators  on  different  wing  sections. 
Through  the  small  angles  of  0°  and  2°  the  low  aspect 
ratios  give  a  maximum  Ky  greater  than  with  the  larger 
aspects.  The  larger  aspects  increase  the  lift  through  a 
larger  range  of  angles  but  have  a  lower  maximum  value 
for  Ky  at  small  angles.  Beyond  2°  the  larger  aspect  ratios 
give  a  greater  Ky. 

Aspect  for  Flat  Plates.  For  flat  plates  the  results  are 
different  than  with  cambered  sections.  The  lift-drag 
ratios  are  not  much  improved  with  an  increase  in  aspect, 
but  the  highest  maximum  lift  is  obtained  with  a  small 


160  WING  SECTIONS 

aspect  ratio.  For  this  reason,  a  small  aspect  ratio  should 
be  used  when  a  high  lift  is  to  be  obtained  at  low  speeds 
with  a  flat  plate  as  in  the  case  of  control  surfaces.  An 
aspect  ratio  of  unity  is  satisfactory  for  flat  vertical  rud- 
ders since  a  maximum  effect  is  desirable  when  taxi-ing 
over  the  ground  at  low  speeds.  The  flat  plate  effects  are 
not  important  except  for  control  surfaces,  and  even  in 
this  case  the  plates  are  being  superseded  by  double  cam- 
bered sections. 

Reason  for  Aspect  Improvement.  The  air  flows  later- 
ally toward  the  Aving  tips  causing  a  very  decided  drop  in 
lift  at  the  outer  ends  of  the  wings.  The  lift-drag  ratio  is 
also  reduced  at  this  point.  The  center  of  pressure  moves 
back  near  the  trailing  edge  as  we  approach  the  tips,  the 
maximum  zone  of  suction  on  the  upper  surface  being  also 
near  the  trailing  edge.  The  lift-drag  ratio  at  the  center  of 
the  plane  is  between  4  or  5  times  that  at  a  point  near  the 
tips.  All  of  the  desirable  characteristics  of  the  wing  are 
exhibited  at  a  point  near  the  center. 

When  the  aspect  ratio  is  increased,  the  inefiicient  tips 
form  a  smaller  percentage  of  the  total  wing  areas,  and 
hence  the  losses  at  the  tips  are  of  less  importance  than 
would  be  the  case  with  a  small  aspect.  The  end  losses 
are  not  reduced  by  end  shields  or  plates,  and  in  attempts 
to  prevent  lateral  flow  by  curtains,  the  losses  are  actually 
often  increased.  Proper  design  of  the  form  of  the  wing 
tip,  such  as  raking  the  tips,  or  washing  out  the  camber  and 
incidence,  can  be  relied  upon  to  increase  the  lift  factor. 
This  change  in  the  tips  causes  the  main  wind  stream  to 
enter  the  wings  in  a  direction  opposite  to  the  lateral  leak- 
age flow  and  therefore  reduces  the  loss.  Properly  raked 
tips  may  increase  the  lift  by  20  per  cent. 

Effects  of  Scale  (Size  and  Velocity).  In  the  chapter 
^'Elementary  Aerodynamics"  it  was  pointed  out  that  the 
lift  of  a  surface  was  obtained  by  the  motion  of  the  air,  or 
the  "turbulence"  caused  by  the  entering  of  the  plane.    It 


WING  SECTIONS  161 

was  also  explained  that  the  effect  of  the  lift  due  to  turbu- 
lence varied  as  the  square  of  the  velocity  and  directly  as 
the  area  of  the  wings.  This  would  indicate  that  the  lift 
of  a  small  wing  (Model)  would  be  in  a  fixed  proportion  to 
a  large  wing  of  the  same  type.  This  holds  true  in  practice 
since  nearly  all  laboratories  have  found  by  experiment 
that  the  lift  of  a  large  wing  could  be  computed  directly 
from  the  results  obtained  with  the  model  without  the  use 
of  correction  factors.  That  is  to  say,  that  the  lift  of  a  large 
wing  with  40  times  the  area  of  the  model,  would  give  40 
times  the  lift  of  the  model  at  the  same  air  speed.  In  the 
same  way,  the  lift  would  be  proportional  to  the  squares 
of  the  velocities.  If  the  span  of  the  model  is  taken  at  ''1" 
feet,  and  the  velocity  as  V^  feet  per  second,  the  product  IV 
would  represent  both  the  model  and  the  full  size  machine. 
The  lift  is  due  to  aerodynamic  forces  strictly,  and  hence 
there  should  be  no  reason  why  the  "V-"  law  should  be 
interfered  with  in  a  change  from  the  model  to  the  full  size 
machine. 

In  the  case  of  drag  the  conditions  are  difterent,  since  » 
the  drag  is  produced  by  two  factors  that  vary  at  diiTerent  ,    ^k^,*^! 
rates.     Part  of  the  drag  is  caused  l)y  turbulence  or  aero-  i  Avw-f /w 
dynamic  forces  and  part  by  skin  friction,  the  former  vary-  llt^^f'^^' '' 
ing  as  V-  while  the  skin  friction  varies  as  V^-^^.    The  aero- 1  iv*^ 
dynamic  drag  varies  directly  with  the  area  or  span  while 
the  skin  friction  part  of  the  drag  varies  as  P-^',  where  1 
is  the  span.     From  considerations  of  the  span  and  the 
speed,  it  will  be  seen  that  the  frictional  resistance  increases 
much  slower   than  the  aerodynamic  resistance,  and  conse-j 
quently  the  large  machine  at  high  speed  would  give  less 
drag  and  a  higher  value  of  L/D  than  the  small  model. j 
In  other  words,  the  results  of  a  model  test  must  be  cor-/ 
rected  for  drag  and  the  lift-drag  ratio  when  applied  to  a 
full  size  machine.    Such  a  correction  factor  is  sometimesj 
known  as  the  "Scale  factor." 

Eiffel  gives  the  correction  factor  as  1.08,  that  is  the  lift- 


162  WING  SECTIONS 

drag  ratio  of  the  full  size  machine  will  be  approximately 
1.08  times  as  great  as  the  model. 

A  series  of  full  size  tests  were  made  by  the  University 
of  St.  Cyr  in  1912-1913  with  the  object  of  comparing  full 
size  aeroplane  wings  with  small  scale  models  of  the  same 
Aving  section.  The  full  size  wings  were  mounted  on  an 
electric  trolley  car  and  the  tests  were  made  in  the  open  air. 
Many  differences  were  noted  when  the  small  reproduc- 
tions of  the  W'ings  were  tested  in  the  wind  tunnel,  and  no 
satisfactory  conclusions  can  be  arrived  at  from  these  tests. 
According  to  the  theory,  and  the  tests  made  by  the  N.  P. 
L.,  the  lift-drag  ratio  should  increase  with  the  size  but 
the  St.  Cyr  tests  showed  that  this  was  not  always  the 
case.  In  at  least  three  of  the  tests,  the  model  showed 
better  results  than  the  full  size  machine.  There  seemed 
to  be  no  fixed  relation  between  the  results  obtained  by  the 
model  and  the  large  wing.  The  center  of  pressure  move- 
ment was  always  different  in  every  comparison  made. 

One  cause  of  such  pronounced  difference  would  prob- 
ably be  explained  by  the  difference  in  the  materials  used 
on  the  model  and  full  size  wing,  the  model  wing  being 
absolutely  smooth  rigid  wood  while  the  full  size  wing  was 
of  the  usual  fabric  construction.  The  fabric  would  be 
likely  to  change  in  form  under  different  conditions  of 
angle  and  speed,  causing  a  great  departure  from  the  true 
values.  Again,  the  model  being  of  small  size,  would  be  a 
difficult  object  to  machine  to  the  exact  outline.  A  differ- 
ence of  1/1000  inch  from  the  true  dimension  would  make 
a  great  difference  in  the  results  obtained  with  a  small 
surface. 

Plan  Form.  Wings  are  made  nearly  rectangular  in 
form,  with  the  ends  more  or  less  rounded,  and  very  little 
is  now  known  about  the  effect  of  wings  varying  from  this 
form.  Raking  the  ends  of  the  wing  tips  at  a  slight  angle 
increases  both  the  lift-drag  and  lift  by  about  20  per  cent, 
the  angle  of  the  raked  end  being  about  15  degrees.    Raking 


WING  SECTIONS  163 

is  a  widely  adopted  practice  in  the  United  States,  espe- 
cially on  large  machines. 

Summary  of  Corrections.  Wt  can  now  work  out  the 
total  correction  to  be  made  on  the  wind  tunnel  tests  for  a 
full  size  machine  of  any  aspect  ratio.  The  lift  co-efficient 
should  be  used  as  given  by  the  model  test  data,  but  the 
corrections  can  be  applied  to  the  lift-drag  ratio  and  the 
drag.  The  scale  factor  is  taken  at  1.08,  the  form  factor 
due  to  rake  is  1.2,  and  the  aspect  correction  is  taken  from 
the  foregoing  table.  The  total  correction  factor  will  be 
the  product  of  all  of  the  individual  factors. 

Example.  A  certain  wing  section  has  a  lift-drag  ratio 
of  15.00,  as  determined  by  a  wind  tunnel  test  on  a  model, 
the  aspect  of  the  test  plane  being  6.  The  full  size  wing  is 
to  have  an  aspect  ratio  of  8,  and  the  wing  tips  are  to  be 
raked.  What  is  the  corrected  lift-drag  ratio  of  the  full 
size  machine  at  14°  ? 

Solution.  The  total  correction  factor  will  be  =1.08X 
1. 10 X  1.2=1.439.  The  litt-drag  ratio  of  the  full  size  modi- 
fied wing  becomes  15.00X1.439=21.585. 

As  a  comparison,  we  will  assume  the  same  wing  section 
with  rectangular  tips  and  an  aspect  ratio  of  3.  The  total 
correction  factor  for  the  new  arrangement  is  now  1.08X 
0.72=0.7776  where  0.72  is  the  relative  lift-drag  due  to  an 
aspect  of  3.  The  total  lift-drag  is  now  15.00X0.7776= 
11.664. 

Having  a  large  aspect  ratio  and  raked  tips  makes  a  very 
considerable  difference  as  will  be  seen  from  the  above 
results,  the  rake  and  aspect  of  8  making  the  difference 
between  21.585  and  11.664  in  the  lift-drag.  Area  for  area, 
the  drag  of  the  first  plane  will  be  approximately  one-half 
of  the  drag  due  to  an  aspect  ratio  of  three. 

Lift  in  Slip  Stream.  The  portions  of  a  monoplane  or 
tractor  biplane  lying  in  the  propeller  slip  stream  are  sub- 
jected to  a  much  higher  wind  velocity  than  the  outlying 
parts  of  the  wing.     Since  the  lift  is  proportional  to  the 


164  WING  SECTIONS 

velocity  squared,  it  will  be  seen  that  the  lift  in  the  slip 
stream  is  far  higher  than  on  the  surrounding  area.  As- 
suming for  example,  that  a  certain  propeller  has  a  slip  of 
30  per  cent  at  a  translational  speed  of  84  miles  per  hour, 
the  relative  velocity  of  the  slip  stream  will  be  84/0.70= 
120  miles  per  hour.  Assuming  a  lift  factor  (Ky)=0.0022, 
the  lift  in  the  slip  stream  will  be  L==0.0022X  120X120= 
31.68  pounds  per  square  foot.  In  the  translational  wind 
stream  of  84  miles  per  hour,  the  lift  becomes  L=0.0022X 
84X84=15.52  pounds  per  square  foot.  In  other  words, 
the  lift  of  the  portion  in  the  slip  stream  is  nearly  double 
that  of  the  rest  of  the  wing  with  a  propeller  efficiency  of 
70  per  cent. 


CHAPTER  VII. 
BIPLANES  AND  TRIPLANES. 

Biplane  Characteristics.  From  an  aerodynamic  stand- 
point, the  monoplane  wing  is  more  efficient  than  the  super- 
posed wings  of  the  biplane  type,  since  the  proximity  of 
the  two  surfaces  in  the  latter  causes  a  decided  loss  in  the 
total  lift.  Other  practical  advantages,  however,  ofifset  the 
losses  due  to  the  superposed  surfaces,  and  hence  the  total 
efficiency  of  the  complete  biplane  may  be  even  greater 
than  that  of  the  monoplane.  For  the  same  area  the  struc- 
tural parts  of  the  biplane  are  lighter,  and  this  advantage 
increases  rapidly  w^ith  the  size  of  the  machine  so  that 
when  a  span  of  36  feet  is  exceeded,  any  other  arrangement 
than  that  of  the  biplane  or  triplane  becomes  almost  a 
practical  impossibility.  A  biplane  is  easier  and  cheaper 
to  make  than  a  monoplane,  since  the  wing  bracing  of  the 
former  can  be  arranged  to  better  advantage,  the  load- 
bearing  members  can  be  simpler,  and  the  safety  factor 
made  higher  for  an  equal  weight.  By  suitable  adjust- 
ments between  the  wings  of  a  biplane,  it  is  possible  to 
obtain  a  very  high  degree  of  inherent  longitudinal  sta- 
bility without  incurring  much  loss  in  efficiency,  an  ar- 
rangement that  is  of  course  impossible  with  a  single 
monoplane  surface.  By  ''staggering,"  the  view  of  the 
pilot  is  increased,  and  the  generally  smaller  size  of  the 
machine  permits  of  better  maneuvering  qualities  for  a 
given  load. 

Interference.  Due  to  "interference,"  or  to  the  choking 
of  the  air  stream  between  the  upper  and  lower  surfaces, 
the  lift  of  both  wings  is  reduced,  with  the  drag  remaining 

165 


166  BIPLANES  AND  TRIPLANES 

about  the  same  as  with  a  single  surface.  This,  of  course, 
reduces  the  total  lift-drag  ratio  at  all  except  certain  angles. 
The  relative  lift-drag  ratios  of  the  monoplane  and  biplane 
depend  to  some  extent  upon  the  form  of  the  wing.  Inter- 
ference causes  a  loss  on  the  opposing  faces  of  the  wings, 
the  lift  being  reduced  on  the  top  surface  of  the  lower  wing, 
and  on  the  bottom  surface  of  the  top  wing.  Since  the 
upper  surface  of  the  lower  wing  is  under  suction,  and 
therefore  produces  the  greater  proportion  of  lift,  it  is  natu- 
ral that  the  lower  wing  lift  should  be  reduced  to  a  greater 
extent  than  in  the  upper  wing,  since  it  is  only  the  lower 
surface  of  the  latter  that  is  affected.  At  normal  flight 
angles  the  upper  wing  carries  apout  55  per  cent  of  the 
total  load.  At  zero  degrees  incidence,  the  upper  wing 
carries  as  high  as  62  per  cent  of  the  total  load,  while  at 
12  degrees  this  may  be  reduced  to  54  per  cent. 

Gap-Chord  Ratio,  Calling  the  distance  between  the 
upper  and  lower  wings  the  **gap,"  it  may  be  said  that  the 
ratio  of  the  gap  to  the  wing  chord  greatly  influences  the 
lift.  This  ratio  is  called  the  "gap-chord  ratio,"  and  may 
vary  from  0.8  to  1.0  in  small  machines  or  1.0  to  1.2  in  slow, 
heavy  aeroplanes.  With  the  drag  remaining  practically 
constant,  the  lift-drag  is  of  course  affected  by  a  change 
in  the  gap-chord  ratio,  this  quantity  being  diminished  at 
small  gap  ratios.  Compared  with  a  monoplane,  the  lift 
of  a  biplane  is  about  0.77  when  the  gap  is  0.8  of  the  chord, 
and  about  0.89  of  the  monoplane  value  when  the  gap-chord 
ratio  is  increased  to  1.6.  In  this  range  the  lift-drag  approx- 
imates 0.82  and  0.89,  respectively.  The  center  of  pressure 
movement  is  not  greatly  changed  with  any  gap-chord 
ratio,  and  to  all  practical  purposes  remains  the  same  as 
with  the  monoplane.  It  should  be  understood  that  these 
remarks  apply  only  to  the  "Orthogonal"  biplane  arrange- 
ment in  which  the  wings  are  vertically  over  one  another. 

While  biplane  efificiency  is  increased  by  having  a  large 
gap-chord  ratio  (wing  efficiency  alone),  the  total  efficiency 


BIPLANES  AND  TRIPLANES  167 

of  the  aeroplane  is  not  always  increased  by  a  large  gap, 
principally  because  of  the  great  head  resistance  due  to 
the  longer  struts  and  interplane  bracing.  At  high  speeds 
the  longer  bracing  members  often  more  than  offset  the 
gain  due  to  wing  efficiency,  and  as  a  result  the  gap  of  high 
speed  scouts  will  generally  be  found  in  the  neighborhood 
of  0.8  the  chord.  With  slow,  heavy  machines,  where  lift 
is  of  great  importance,  and  where  slow  speed  does  not 
aft'ect  the  structural  resistance  to  so  great  an  extent,  the 
gap-chord  ratio  will  range  from  1.0  to  1.2. 

In  making  the  above  comparisons  between  monoplanes 
and  biplanes,  equal  aspect  ratios  have  been  assumed  for 
both  types,  but  in  actual  practice  the  aspect  ratio  of 
biplanes  is  always  greater  than  with  monoplanes,  and  as  a 
result  the  biplane  loss  is  usually  less  than  indicated  above. 
When  correction  has  been  made  for  the  aspect  ratio,  the 
disparity  in  the  monoplane  and  biplane  values  of  Ky  and 
L/D  is  not  as  great  as  commonly  supposed.  ''Biplane 
reduction  factors,"  or  the  factors  used  in  reducing  mono- 
plane values  to  those  of  the  biplane,  depend  to  a  great 
extent  upon  the  wing  section  as  well  as  upon  the  gap,  and 
for  exact  values  of  the  factors  we  should  have  the  tests 
report  of  the  wings  in  biplane  form.  Lacking  this  infor- 
mation, we  can  adopt  the  values  obtained  by  the  N.  P.  L. 
for  an  old  type  of  wing  in  order  to  get  approximate  results. 
To  obtain  the  biplane  values,  multiply  the  monoplane 
values  obtained  by  the  wind  tunnel  test  by  the  factors 
found  under  the  required  gap-chord  ratio.  These  factors 
apply  to  an  aspect  ratio  of  6. 

BIPLANE   REDUCTION    FACTORS    (n.    P.    L.) 

(At  Normal  Flight  Angles) 

Gap-Chord  Ratio 0.8  1.0  1.2  1.6 

Ky  Reduction  Factor 0.77        0.82        0.86        0.89 

L/D  Reduction  Factor 0.82        0.84        0.85         0.89 


168 


BIPLANES  AND  TRIPLANES 


Dr.  Hunsaker  conducted  experiments  at  the  Massachu- 
setts Institute  of  Technology  on  biplane  and  triplane  com- 
binations, and  the  results  were  reported  in  ''Aviation  and 
Aeronautical  Engineering,"  Nov.  1,  1916.  The  R.  A.  F.-6 
section  was  used  with  a  gap-chord  ratio  of  1.2.  The 
biplane  portions  of  the  experiments  are  as  follows,  the 
actual  Ky  and  L/D  values  and  reduction  factors  being 
arranged  according  to  the  angle  of  incidence : 


Biplane   reduction   factors    (aspect   ratio   =   6.     Gap   chord 


1.2.) 


ANGLE  OF 

LIFT  COEFFICIENT  KJ 

LIFT DRAG 

RATIO  L/D 

INCIDENCB 

ACTUAL  Ky 

FACTOR 

ACTUAL 

FACTOR 

0° 

0.000432 

0.888 

6.30 

0.732 

2° 

.000864 

.838 

12.20 

0.747 

4° 

.001230 

.854 

13.80 

0.820 

8° 

.001860 

.852 

11.30 

0.819 

12° 

.002440 

.876 

9.50 

0.950 

le'' 

.002730 

.985 

5.60 

1.240 

It  will  be  noted  that  there  is  steady  improvement  in  the 
lift  factor  with  an  increase  in  the  angle  from  2°  up  (except 
at  8°),  and  that  the  same  holds  true  with  the  L/D  factor. 
That  is,  the  biplane  values  become  nearly  monoplane 
values  at  high  angles,  and  in  the  case  of  the  L/D  ratio 
the  biplane  actually  is  24  per  cent  greater  than  the  mono- 
plane value  at  an  angle  of  16°.  The  lift  coefficient  Ky, 
above,  is  not  far  from  the  corresponding  Ky,  for  gap- 
chord  ratio  =  1.2  in  the  first  table.  The  maximum  biplane 
value  of  L/D  occurs  at  the  same  point  as  in  the  monoplane 
wing,  that  is,  at  4°.  The  fact  that  the  lift-drag  is  so  high 
at  16°  is  very  favorable,  since  the  biplane  would  be  less 
likely  to  stall  when  flying  slowly,  and  with  a  big  demand 
on  the  engine.  The  range  of  angles  at  the  stalling  angle 
is  much  greater  than  with  the  monoplane  wing,  and  the 
lift  does  not  fall  off  so  rapidly  after  the  maximum  is 
reached. 

Biplane  Arrangements.  In  the  foregoing  data  we  have 
assumed  that  the  upper  wing  was  placed  directly  above 


BIPLANES  AND  TRIPLANES 


169 


the  lower,  and  with  the  leading  edges  on  the  same  vertical 
line  as  shown  by  Fig.  3.  This  is  known  as  an  "Orthog- 
onal" biplane,  and  the  gap  is  indicated  by  G  and  the  chord 
by  C.  In  Fig.  4  the  forward  edge  of  the  top  wing  is 
advanced  beyond  the  lower,  or  is  ''Staggered,"  the  amount 
of  the  stagger  being  indicated  by  S.    This  allows  of  better 


Figs.   3-8.     Different  Biplane  Arrangements,  Showing  Stagger  and  Decalage. 

view,  and  slightly  increases  both  the  lift  and  L/D  values. 
With  a  comparatively  large  stagger  the  range  of  the 
stalling  angle  is  increased,  and  the  Hft  does  not  fall  ofif 
as  rapidly  after  the  maximium  is  reached  as  with  the 
orthogonal  type.  In  Fig.  5  the  top  wing  is  given  a  back- 
ward stagger,  but  the  exact  effects  of  this  arrangement 
are  not  generally  known.    There  are  few  machines  using 


170  BIPLANES  AND  TRIPLANES 

the  reversed  stagger,  the  only  example,  to  the  writer's 
knowledge,  being  the  De  Havilland  speed  scout.  By 
staggering,  the  resistance  of  the  interplane  bracing  struts 
(3)  is  somewhat  reduced,  because  of  their  inclination  with 
the  wind,  although  they  are  longer  for  the  same  gap  than 
in  Fig.  3. 

Fig.  6  shows  the  chord  of  the  lower  wing  (C)  shorter 
than  the  upper  chord,  a  type  used  in  the  Nieuport  speed 
scout.  In  effect,  this  is  a  form  of  stagger,  and  it  undoubt- 
edly widens  the  view  of  the  pilot,  and  to  some  extent 
increases  the  efficiency  and  the  range  of  the  stalling  angle. 
Neither  the  stagger  in  (4)  nor  the  small  lower  chord  alone 
improves  the  stability  to  any  extent.  To  obtain  any 
marked  advantage  with  the  short  lower  chord,  the  chord 
C  must  be  very  much  shorter  than  the  upper  chord,  say 
from  0.80C  to  0.50C.  The  loss  of  area  is  so  great  that 
this  would  not  be  permissible  on  any  except  the  fastest 
machines,  where  lift  is  not  a  primary  consideration.  The 
pilot's  view,  however,  is  very  much  improved  with  the 
short  lower  chord,  and  in  battle  this  is  an  important 
consideration. 

Fig.  7  shows  the  chord  of  the  upper  wing  inclined  at 
an  angle  with  the  lower  chord  by  the  amount  (d).  This 
is  known  as  "Decalage"  and  is  productive  of  a  great 
degree  of  longitudinal  stability  when  taken  in  combina- 
tion with  stagger.  The  stability  attained  by  decalage  and 
stagger  is  without  a  great  loss  in  the  L/D  ratio,  while  the 
lift  and  stalling  angle  range  are  both  increased.  This 
latter  stable  combination  is  shown  by  Fig.  8,  in  which  the 
wings  are  given  both  stagger  and  decalage. 

Forward  Stagger.  Eiffel  performed  experiments  with 
Dorand  wings,  and  found  that  when  the  top  surface  was 
staggered  forward  by  1/2.5  of  the  chord  (0.4C),  and  with  a 
gap-chord  ratio  of  0.9,  an  Increase  in  lift  of  from  6  to  10 
per  cent  was  obtained.  The  L/D  was  the  same  as  with 
no  stagger.  With  thin  circular  plates,  1/13.5  camber,  and 


BIPLANES  AND  TRIPLANES 


171 


a  gap-chord  ratio  =  0.66,  the  lift-drag  was  better  (than 
with  no  stagger)  only  when  the  value  of  Ky  was  greater 
than  0.066  (metric).     Then  the  L/D  improved  progres- 


Slow  Speed,  Two-Seat  Biplane,  with  a  Large  Gap-Chord  Ratio.  The  Large 
Gap  Is  Permissible  in  a  Slow  Machine,  as  the  Strut  Resistance  Is 
Less  Than  the  Gain  in  Lift-Drag  Ratio  Obtained  by  the  Greater 
Gap.  It  Will  Be  Noted  That  These  Wings  Have  a  Considerable 
Amount  of  Stagger.  The  Position  of  the  Bottom  Wing  Allows  the 
Observer  to  See  Almost  Directly  Below. 


A  High  Speed,  Two-Seat  Fighting  Biplane,  with  a  Small  Gap-Chord  Ratio. 
In  This  Case,  the  Strut  Resistance  Would  Be  Greater  Than  the 
Aerodynamic  Gain  of  the  Wings  with  a  Greater  Gap  Chord  Ratio. 
The  Gunner  Is  Located  in  the  Rear  Seat,  and  Behind  the  Trailing 
Edge  of  the  Lower  Wings.  He  Has  a  Clear  Field  to  the  Rear  and 
Over  the  Top  Wing. 


sively  with  the  amount  of  stagger.  Ky  was  improved  by 
5  per  cent  when  the  stagger  was  equal  to  half  the  chord, 
and  by  10  per  cent  when  the  stagger  was  equal  to  the 


172  BIPLANES  AND  TRIPLANES 

chord.  The  N.  P.  L.  with  a  Bleriot  wing,  aspect  ratio=4, 
found  that  Ky  was  increased  by  5  to  6  per  cent  with  a 
stagger  of  0.4C,  and  the  L/D  was  increased  by  about  4 
percent.     The  gap-chord  ratio  was   1.00. 

In  a  series  of  tests  made  by  A.  Tcherschersky,  the  back- 
ward stagger  as  in  Fig.  5  gave  about  15  per  cent  greater 
lift  than  the  orthogonal  biplane,  or  about  4  per  cent  less 
lift  than  a  monoplane  surface  of  the  same  area.  The 
stagger  in  this  experiment  was  about  0.33C.     In  default 


A  Single  Seat  Biplane  Speed  Scout  with  an  Air  Cooled  Motor. 

of  more  accurate  information,  it  would  seem  that  back- 
ward stagger  would  give  better  results  than  forward  stag- 
ger, since  the  air  swept  down  by  the  upper  surface  would 
pass  further  to  the  rear  of  the  lower  plane  and  hence 
would  not  so  greatly  affect  the  vacuum  on  the  upper  sur- 
face of  the  lower  wing.  This  would,  however,  destroy  the 
view  of  the  pilot  to  a  greater  extent  than  any  of  the  other 
arrangements. 

Stagger  always  introduces  structural  difficulties,  makes 
the  wings  difficult  to  assemble,  and  the  wires  are  of  vary- 
ing lengths.  A  simple  orthogonal  cell  is  more  compact 
and  better  from  a  manufacturing  standpoint,  as  it  sim- 
plifies the  fittings,  and  to  a  slight  extent  decreases  the 


BIPLANES  AND  TRIPLANES  173 

weight.  When  combined  with  sweep  back,  the  complica- 
tion is  particularly  in  evidence.  It  is  pleasing  to  note  the 
prevalence  of  orthogonal  cells  on  modern  battle-planes. 

Influence  of  Camber.  The  amount  of  air  swept  down 
by  the  upper  wing  is  largely  determined  by  the  curvature 
of  the  under  surface  of  the  upper  wing.  By  decreasing, 
or  flattening  out  the  curvature  of  this  surface,  the  velocity 
is  increased  in  a  horizontal  direction  and  reduced  in  a 
vertical  direction,  so  that  the  lower  wing  is  less  affected. 
The  upper  surface  of  the  upper  wing  is  not  influenced  by 
interference.  It  should  be  noted  at  this  point  that  air  in 
striking  a  convex  surface  is  increased  in  horizontal  speed 
while  the  reverse  is  true  of  the  lower  concave  surface.  If 
the  under  surface  of  the  upper  wing  were  made  convex, 
the  down  trend  of  the  air  would  be  still  further  reduced, 
and  the  loss  on  the  lower  wing  reduced  in  proportion. 

Increasing  the  camber  on  the  upper  surface  of  the  lower 
wing  increases  its  horizontal  velocity  and  hence  affects  the 
upper  wing  to  a  less  extent,  but  as  the  upper  wing  loss  is 
comparatively  slight,  the  camber  increase  below  is  not  of 
great  consequence.  This  has  only  been  tried  in  one  ma- 
chine to  the  writer's  knowledge,  one  of  the  Standard  sea- 
planes, in  which  the  upper  wing  was  an  R.  A.  F.-6  and  the 
lower  wing  was  a  deeply  cambered  U.  S.  A.-2  section. 
The  lower  surface  of  the  R.  A.  F.-6  is  comparatively  flat. 

Effects  of  Decalage.  When  the  upper  wing  incidence  is 
increased  in  regard  to  that  of  the  lower  wing,  or  is  given 
decalage,  the  stability  is  increased  with  a  slight  increase  in 
the  power  or  drag.  This  angle  shown  by  (d)  in  Figs.  7 
and  8,  must  be  accompanied  by  stagger  to  obtain  stability, 
the  angle  (d)  ranging  from  V  to  4°.  With  a  decalage  of 
2.5°,  and  a  stagger  of  half  the  chord,  a  high  degree  of 
stability  is  attained  with  a  loss  in  the  lift-drag  of  from  4  to 
6  percent.  The  lift  and  the  range  of  the  stalling  angle  are 
both  increased,  the  former  by  about  3  percent,  while  the 
latter  is  nearly  double.    By  increasing  the  decalage  to  4°, 


174  BIPLANES  AND  TRIPLANES 

the  lift-drag  is  still  4  percent  less  than  with  the  orthog- 
onal cell,  but  the  range  of  the  stalling  angle  is  nearly 
tripled.  The  4°  decalage  is  very  stable  and  is  suitable  for 
training  machines  or  for  amateurs.  In  either  case,  the 
stagger-decalage  system  is  usually  better  than  sweep  back, 
reflex  curves,  or  negative  wing  tips. 

Without  regard  to  the  stability,  and  only  with  the  idea 
of  a  greater  L/D  in  mind,  it  has  been  usual  in  several 
European  machines  to  adopt  a  ''negative"  decalage ;  that 
is,  to  increase  the  angle  of  the  lower  wing  in  regard  to  the 
upper  chord.  With  the  top  chord  horizontal,  a  negative 
decalage  of  4°  would  make  the  incidence  of  the  lower  wing 
equal  to  4°.  This  has  not  been  generally  found  advantage- 
ous in  model  tests,  but  in  full  size  machines  there  is  a  con- 
siderable increase  in  the  L/D  ratio.  The  greater  incidence 
of  the  lower  wing  also  improves  the  lift  of  this  surface  and 
thus  requires  less  surface  for  obtaining  the  same  total  lift, 
especially  when  top  wing  is  staggered  forward.  Incidence 
of  top  wing  of  Nieuport  =\°-30\  Lower  wing  is  set 
at  3°. 

Varying  Incidence.  \\'ith  several  types  of  European 
speed  scouts,  and  in  the  case  of  the  old  Handley-Page 
monoplane,  the  angle  of  incidence  is  reduced  from  the 
center  of  the  wing  to  the  tip.  Thus  in  one  speed  scout,  the 
incidence  at  the  body  is  4°,  and  2°  at  the  tips.  A  decrease 
in  angle  toward  the  tips  has  much  the  same  effect  as  an 
increase  in  aspect  ratio;  that  is,  it  decreases  the  lateral 
flow  and  end  leakage.  It  also  has  an  efifect  in  aiding  the 
lateral  stability  because  there  is  less  lift  at  the  tips,  and 
hence  they  are  less  aiTected  by  side  gusts.  "Washed 
out"  incidence  is  an  aid  to  longitudinal  stability,  as  the 
center  of  pressure  at  the  tips  is  moved  further  back  than 
at  the  center  of  the  wing,  and  therefore  the  C.  P.  is  dis- 
tributed over  a  longer  distance  fore  and  aft  than  it  would 
be  with  a  uniform  angle  of  incidence. 

In  driving  the  propeller,  the  motor  tends  to  turn  the 


BIPLANES  AND  TRIPLANES  175 

body  in  a  direction  opposite  to  that  of  the  propeller  rota- 
tion, and  if  no  other  provision  is  made  this  must  be  over- 
come by  means  of  the  ailerons.  The  ''Motor  torque"  on 
small  span  machines  is  particularly  difficult  to  overcome  in 
this  way,  owing  to  the  short  lever  arm  length  of  the  ailer- 
ons. To  practically  overcome  the  torque,  without  exces- 
sively loading  the  ailerons,  it  is  usually  the  practice  to  set 
the  lower  left  wing  tip  at  a  greater  angle  than  the  lower 
right  wing.  The  greater  angle  at  the  left  gives  a  lift  that 
opposes  the  turning  moment  of  the  motor.  This  com- 
pensation can  never  be  complete,  for  the  motor  torque 
varies  with  the  motor  output,  hence  an  average  angle  is 
selected  so  that  the  incidence  will  cover  the  usual  hori- 
zontal flight  speeds. 

Triplane  Arrangement.  When  a  biplane  exceeds  a  cer- 
tain weight  the  area  required  for  a  given  landing  speed 
makes  it  desirable  to  increase  the  number  of  lifting  sur- 
faces to  more  than  two,  if  the  span  and  stress  are  to  be 
kept  down  within  reasonable  limits.  Thus  the  biplane  has 
its  limits  as  well  as  the  monoplane,  and  in  the  biplane 
this  limit  is  generally  reached  when  the  span  approaches 
80  feet.  In  addition  to  the  increased  weight  due  to  spans 
of  over  80  feet,  there  are  other  troubles  in  regard  to  the 
space  required  for  housing,  and  awkwardness  in  ma- 
neuvering. On  the  smaller  and  faster  aeroplanes,  the 
triplane  arrangement  permits  of  space  condensation,  and 
also  allows  of  larger  aspect  ratios  than  with  the  biplane. 
The  greater  depth  of  the  triplane  structure  makes  the 
interplane  bracing  even  more  effective  than  in  the  case 
of  the  biplane.  For  equal  spans  there  is  less  bracing  ex- 
posed to  the  wind,  and  the  weight  of  the  wing  spars  and 
ribs  can  be  considerably  reduced.  The  shorter  ribs  of  the 
triplane  alone  contribute  in  no  small  degree  to  the  saving 
in  weight. 

Considering  the  wings  alone,  without  reference  to  the 
head  resistance  of  the  bracing,  etc.,  there  is  a  greater  loss 


176  BIPLANES  AND  TRIPLANES 

of  lift  and  L/D  when  three  tiers  of  wings  are  superposed 
than  with  a  biplane.  In  experiments  by  Dr.  Hunsaker 
upon  R.  A.  F.-6  and  Curtiss  wing  sections,  it  was  found 
that  at  about  4°,  that  the  triplane  required  about  6  percent 
more  power  than  the  corresponding  biplane.  At  this 
angle,  the  L/D  for  the  triplane  was  12.8,  against  the  ratio 
of  13.8  for  the  biplane.  The  gap-chord  ratio  in  each  case 
was  maintained  at  1.2.  Both  the  R.  A.  F.-6  and  the  Curtiss 
wings  gave  results  of  the  same  general  character,  and 
there  was  not  a  great  deal  of  difference  in  the  numerical 
values.  At  very  high  angles,  12°  to  16°,  the  lift  of  the 
biplane  and  triplane  only  differed  by  about  2  percent, 
but  at  very  small  angles  such  as  are  used  at  normal  flight 
speeds,  the  reduction  of  lift  in  the  triplane  was  very 
marked. 

The  drag  was  not  greatly  different  below  12°,  but  at  16° 
the  drag-coefficient  is  less  than  that  of  either  the  biplane 
or  monoplane,   and   for  machines  flying  at  low   speeds, 
or   heavily  loaded,   this   decrease  is  of  great  advantage 
since  it  relieves  the  motor  at  a  time  when  power  is  par- 
ticularly required.     At  this  point  it  should  be  noted  that 
at  high  angles,  the  L/D  generally  is  better  for  multiplanes^ 
in  an  almost  direct  proportion  to  the  number  of  surfaces. 
In  this  experiment,  the  lift-drag  ratios  for  a  monoplane, 
biplane,  and  triplane  were  respectively  4.5,  5.6,  and  6.5. 
The  drop  in  lift  after  the  point  of  maximum  lift,  or  the 
stalling  angle,  is  not  as  rapid  as  in  the  case  of  the  biplane 
or  monoplane,  and  hence  there  is  less  danger  of  stalling 
the  triplane.  With  the  same  area,  and  loading,  the  landing 
speed  of  the  biplane  and  triplane  will  be  about  the  same. 
The  following  tables  give  the  lift,  and  lift-drag  ratios 
as  determined  in  these  experiments,  the  factors  being  in 
terms  of  the  monoplane  values  of  an  R.  A.  F.-6  wing.  Thus 
to  obtain  triplane  values,  multiply  the  given  monoplane 
values  by  that  number  opposite   the  required   angle  of 
incidence.     Aspect  ratio  =  6. 


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BIPLANES  AND  TRIPLANES 


MONOPLANE  VALUES 

TUIPLANE  LIFT 

TUIPLANE  L/D 

ANGLE  OF 

fKY) 

INCIDENCE 

Ky  ACTUAL 

L/D  ACTU.\L 

FACTOR 

ACTUAL 

0.830 

L/D  ACTUAL 

FACTOR 

0 

0.000486 

8.60 

0.000404 

6.1 

0.708 

2 

.00103 

16.30 

0.000776 

0.754 

11.4 

.698 

4 

.00145 

16.80 

0.001097 

.757 

12.8 

.761 

8 

.00218 

13.80 

.001690 

.774 

11.1 

.804 

12 

.00278 

10.00 

.002260 

.812 

8.9 

.890 

16 

.00277 

4.5 

.00267 

.964 

6.5 

1.450 

Thus,  if  the  monoplane  lift  value  for  the  R.  A.  F. — 6 
wing  at  4°  is  Ky  =  0.00145,  then  the  triplane  value  will  be 
0.00145  +  0.757  =  0.001097  as  given  in  the  table.  The 
monoplane  lift-coefficient  of  any  other  wing  section  can 
be  handled  in  the  same  way  with  fair  accuracy.  To  obtain 
the  corrected  lift-drag  ratio  for  any  wing  section,  multiply 
the  lift-drag  of  the  monoplane  wing  by  the  factor  in  the 
above  table  corresponding  to  the  incidence  of  the  mono- 
plane test  wing. 

The  upper  wing  gives  the  greatest  percentage  of  lift, 
and  the  middle  wing  the  least,  since  the  latter  suffers  from 
interference  on  both  sides.  It  has  been  found  that  the 
sum  of  the  top  and  bottom  wings  of  a  triplane  group 
gives  the  same  lift  as  the  two  wings  of  a  biplane  under 
equal  conditions.  It  was  also  found  that  the  lift-coeffi- 
cients and  lift-drag  of  the  upper  plane  alone  was  very 
nearly  equal  to  the  lift  of  the  combined  effects  of  all 
three  wrings,  and  at  all  angles.  Calling  the  lift  of  the 
middle  wing  1.00  (4°),  the  lift  of  the  upper  wing  will  be 
1.91  and  the  lower  wing  1.64.  Calling  the  L/D  of  the 
middle  wing  LOO  (4°),  the  relative  life-drag  will  be  L/D 
=  2.59  for  the  upper  wing  and  1.69  for  the  lower.  With  the 
middle  w^ing  still  assumed  at  unity,  the  lift  of  the  top 
plane  is  at  1.49  at  16°,  and  the  lower  wing  1.20.  The  lift- 
drag  at  16  degrees  will  be  respectively  1.00,  1.22,  and 
1.117  for  middle  top  and  bottom.  At  0°,  the  upper  wing 
will  carry  2.68,  the  middle  1.00,  and  the  bottom  1.82.  At 
0°,  the  lift-drag  of  the  top  is  3.63,  the  middle  1.00,  and 


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180  BIPLANES  AND  TRIPLANES 

bottom  2.30.  These  relative  figures  are  only  useful  in 
comparing  the  loading  when  computing  the  strength  of 
the  structural  parts.  See  "Aviation  and  Aeronautical  En- 
gineering" Nov.  1,  1916. 

Overhanging  Wing  Tips.  In  many  American  machines, 
and  in  some  European  machines,  such  as  the  Farman, 
the  upper  wing  is  given  a  much  greater  span  than  the 
lower.  Of  late,  the  tendency  has  been  to  make  the  wings 
of  equal,  span  and  fully  90  per  cent  of  the  modern  ma- 
chines will  be  found  to  be  arranged  in  this  way.  While 
the  overhanging  tips  may  slightly  increase  the  efficiency 
of  the  biplane  by  reducing  interference  at  the  ends,  it 
makes  the  span  unduly  long  and  difficult  to  brace  at 
the  end.  The  added  end  bracing  due  to  the  overhang 
probably  offsets  any  aerodynamic  advantage  to  be  ob- 
tained, although  I  have  no  accurate  data  on  this  point. 
Compactness  is  certainly  not  a  feature.  It  is  said  that 
ailerons  are  more  effective  when  mounted  on  the  upper 
overhang,  and  this  may  be  so,  but  I  note  that  the  area 
is  about  the  same  in  any  case.  With  overhanging  tips, 
the  ailerons  are  generally  placed  on  upper  wings,  only 
while  with  equal  or  nearly  equal  spans,  they  are  placed 
top  and  bottom.  The  overhanging  section  and  the  ailer- 
ons form  a  single  detachable  unit  as  a  general  rule.  With 
nearly  equal  spans,  the  upper  and  lower  ailerons  are 
generally  interconnected  with  a  small  strut  in  such  a 
way  that  they  act  together. 

Small  speed  scouts,  rarely  if  ever,  have  any  overhang 
since  the  object  of  these  machines  is  to  make  them  as 
small  and  compact  as  possible. 


CHAPTER  A  III. 

EFFECTS  OF  PLAN  FORM. 
(TANDEM  AEROPLANES.) 

General  Notes.  Up  to  the  present  we  have  considered 
only  two  wing  outlines,  the  rectangular  and  the  wing 
with  raked  tips.  In  addition,  we  have  considered  only 
the  effect  taking  place  on  monoplane  surfaces.  For  the 
purpose  of  obtaining  longitudinal  stability,  or  for  dis- 
tributing the  lift  upon  two  or  more  following  surfaces, 
the  plan  view  in  some  aeroplanes  has  been  somewhat 
modified  as  in  the  Dunne,  Langley  and  Ago.  Undoubtedly 
the  simple  rectangular  wing,  or  the  wing  with  raked  tips, 
have  proved  the  most  efficient  from  an  aerodynamic 
standpoint,  but  as  the  layman  is  usually  interested  in  dis- 
torted wing  shapes,  or  odd-looking  outlines,  I  will  describe 
the  effect  of  such  stabilizing  forms  upon  the  lift  and  drag. 

Figs.  1-9  show  the  usual  range  of  wing  forms,  at  least 
those  that  have  been  used  on  well  known  machines,  and 
all  of  them  have  flown  with  varying  results.  In  any  case, 
the  variations  in  the  values  of  the  lift  and  lift-drag  are 
not  excessive,  the  extreme  cases  varying  possibly  not 
more  than  20  per  cent  on  either  side  of  the  values  for  a 
plain  rectangular  wing.  While  almost  perfect  longitu- 
dinal stability  can  be  obtained  in  other  ways,  by  less  loss 
than  by  changing  the  plan  form,  certain  manufacturers 
still  adhere  to  one  or  more  deviations  from  the  more  usual 
rectangular  form. 

Fig.  1  shows  a  plan  view  of  a  machine  with  a  rectangu- 
lar wing,  and  No.  2  shows  the  machine  provided  with 
raked  tips.    Fig.  3  is  a  wing  with  an  inclined  entering 

181 


182  PLAN  FORM 

edge,  as  used  on  the  English  Mann  biplane.  Fig.  4  is 
the  German  Ago  with  a  diamond  form  surface,  the  evident 
purpose  of  which  is  to  simplify  the  wing  spar  bracing  as 
shown  by  the  dotted  lines.  By  bringing  the  wing  spars 
together  at  the  tips,  the  spars  themselves  form  a  triangle 
to  resist  the  drag  stresses.  Fig.  5  is  the  common  form  of 
"sweep  back"  or  "retreated  wing"  as  used  in  the  Standard 
H-3  training  biplane,  and  several  other  modern  biplanes. 
While  this  arrangement  undoubtedly  assists  longitudinal 
stability,  it  causes  certain  losses  that  will  be  described 
later.  The  inherently  stable  Dunne  is  shown  by  Fig.  6 
in  which  the  sweep  back  is  increased  to  almost  90  degrees. 
In  fact  the  retreat  is  so  great  that  no  tail  or  stabilizing 
surfaces  are  used  at  all,  the  elevator  functions  being  per- 
formed by  the  ailerons.  It  should  be  noted  that  the  swept 
back  tips  really  act  as  stabilizers  since  they  trail  back  far 
behind  the  center  of  gravity  and  center  of  pressure.  The 
ailerons  (a)  are  not  needed  for  lateral  balance,  hence 
ascent  and  descent — and  also  turning  in  a  horizontal 
plane — are  performed  by  the  ailerons  in  setting  them  in 
different  relative  positions.  Being  far  to  the  rear  they 
are  very  effective  elevators,  although  their  action  and  the 
extreme  retreat  of  the  wings,  causes  a  considerable  drag. 
This  is  somewhat  offset  by  the  absence  of  tail  resistance. 
An  Austrian  or  German  "Taube"  is  shown  by  Fig.  7 
with  the  negative  trailing  wing  tips  (a),  that  greatly  as- 
sist longitudinal  stability,  but  which  are  decidedly  in- 
efficient. This  is  evidenced  by  the  fact  that  neither  the 
Germans  nor  Austrians  build  this  machine  at  present — at 
least  for  active  service.  The  tips  (a)  are  bent  up  at  the 
rear  and  thus  form  a  ''negative"  angle  of  incidence  with 
the  main  lifting  surface.  This  was  the  original  invention 
of  Igo  Etrich,  an  Austrian,  and  as  with  everything  else, 
the  idea  was  promptly  grabbed  by  the  Germans  at  the 
beginning  of  the  war  and  claimed  as  their  own  idea. 
Etrich   was   one  of  the   pioneers   in  aviation,   a   science 


PLAX  FORM 


183 


that  did  not  prosper  in  Germany  until  the  impractica- 
bility of  the  Zeppelin  for  universal  service  was  an  accepted 
fact. 

Fig.  8  is  the  wing  outline  of  the  Bleriot  monoplane,  a 
representative  wing  used  on  the  earlier  monoplanes.  This 
is  really  reversed  rake,  and  hence  does  not  stand  for  effi- 
ciency in  lift.    A  tandem  aeroplane  is  shown  by  Fig.  9 


Figs.    1-9.     Plan  Views  of  Different  Wing  Arrangements  and  Wing  Outlines. 


in  which  the  leading  surface  is  (m)  and  the  trailing  wing 
is  (n).  The  tail  (t)  may,  or  may  not  be  included.  This  is 
a  type  that  has  been  neglected  in  its  practical  development 
although  it  has  been  repeatedly  proposed.  The  Langley 
machine,  the  Montgomery  glider,  and  the  Richardson  are 
of  this  type. 

In  all  the  figures  the  wing  ribs  are  indicated  by  the  thin 
full  lines  passing  across  the  width  of  the  w^ing,  with  the 


184  PLAN  FORM 

tail  at  (t).  The  arrow  represents  the  line  of  flight.  It 
will  be  noted  that  with  any  but  rectangular  wing  outlines, 
the  rib  lengths  are  different,  throughout  the  wing.  This 
makes  this  wing  a  bad  manufacturing  proposition,  and  a 
difficult  and  expensive  wing  to  repair.  To  provide  against 
emergencies  the  aviator  must  keep  a  complete  extra  wing 
in  reserve  for  repair  parts,  while  the  manufacturer  is  put 
to  the  expense  of  a  great  number  of  rib  molds,  and  must 
also  keep  a  large  number  of  ribs  in  stock.  There  is  a  con- 
stant difficulty  due  to  the  fact  that  mistakes  are  often 
made  in  ordering  the  ribs  for  repair,  and  altogether,  any- 
thing but  a  rectangular  wing  is  a  decided  nuisance. 

Sweep  Back  or  Retreat.  With  swept  back  wings,  as 
shown  in  Fig.  5,  the  center  of  pressure  movement  is  pe- 
culiar. The  C.  P.  moves  forward  when  reducing  the  inci- 
dence at  small  angles,  and  thus  tends  to  reduce  head  div- 
ing, but  at  very  large  angles  the  C.  P.  again  moves  for- 
ward, tending  to  increase  the  angles  further  and  thus 
stall  the  machine.  This  reversal  of  C.  P.  movement  takes 
place  at  about  10°  to  12°,  and  the  movement  is  sharper 
and  further  with  each  increase  in  the  sweep  back.  At  ordi- 
nary angles  of  flight,  say  at  from  0°  to  6°  the  forward  C.  P. 
movement  is  satisfactory,  but  at  low  speeds  and  high 
angles  stability  is  only  partially  secured,  and  hence  for 
the  total  performance  sweep  back  is  not  to  be  desired. 
The  wing  section  used  in  the  above  investigation  was  an 
R.  A.  F.  — 6  with  an  aspect  ratio  of  6. 

In  regard  to  the  lift  coefficient  Ky,  it  was  noted  that  this 
factor  was  decreased  with  every  degree  in  the  angle  of 
sweep  back.  At  the  incidence  angle  of  4°  for  the  maximum 
value  of  L/D,  the  value  of  Ky  decreased  from  0.00143 
with  a  sweep  back  of  0°  (Straight  wings),  to  0.00120  with 
a  sweep  back  of  30°.  At  the  same  incidence,  but  with  a 
sweep  back  of  10°,  the  lift  became:  Ky  =  0.00130.  With  a 
retreat  of  20°,  Kyi=  0.00129.  The  lift-drag  also  suffered 
with  an  increase  in  retreat,  this  being  17.00  with  straight 


PLAN  FORM 


185 


wings,  at  16.5  at  10°,  16.2  at  20°,  and  12.8  at  30°.  Up  to, 
and  including  a  retreat  of  20°,  the  loss  in  lift-drag  is  not 
so  bad,  but  in  the  change  from  20°  to  30°  there  is  a  very 
great  loss. 

The  "angle  of  retreat"  herein  specified  is  such  that  each 
wing  is  moved  back  through  an  angle  of  one-half  the  total 


^ 

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w 

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. 

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Si 

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ton* 

VII 

tmr 

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Fig.    10.      ^ Upper)    Various    Angles    Made    by    Wings    During    Experiments. 
(Below)  The  Center  of  Pressure  Movement  with  Varying  Angles. 


given  retreat  (r)  angle  in  Fig.  5.  That  is,  with  a  retreat 
of  30°,  each  wing  section  makes  an  angle  of  15°  with  the 
entering  edge  of  a  pair  of  straight  wings.  It  is  more  usual 
to  specify  the  included  angle  between  the  leading  edges 
as  indicated  by  (S)  in  Fig.  5.  In  the  upper  portion 
of  Fig.  10  is  shown  the  various  angles  of  the  wings  during 
the  experiments,  while  below  is  the  C.  P.  movement  ac- 


186  PLAN  FORM 

cording  to  the  different  angles  of  retreat.  The  above  is 
based  on  experiments  made  by  H.  E.  Rossell  and  C.  L. 
Brand,  assistant  Naval  Constructors,  U.  S.  Navy,  and 
published  in  "Aerial  Age." 

The  center  of  pressure  referred  to  is  that  at  the  forward 
point  of  the  middle  longitudinal  section  of  each  wing,  and 
with  a  given  incidence  the  C.  P.  is  thrown  to  the  rear  by 
about  0.2  of  the  chord  by  a  sweep  back  of  10  degrees,  and 
0.4  of  the  chord  for  a  sweep  back  of  20°.  The  center  of 
gravity  of  the  machine  will  thus  have  to  be  moved  to  the 
rear  if  sweep  back  is  employed. 

In  making  a  turn  a  machine  with  sweep  back  has  a  nat- 
ural tendency  to  bank  up  in  the  correct  attitude,  and  even 
a  retreat  of  10°  will  add  nearly  100  per  cent  to  the  banking 
tendency  when  compared  to  a  pair  of  straight  wings.  It 
will  be  seen  that  with  swept  back  wings,  the  leading  edge 
is  radial,  consequently  meets  the  air  stream  at  more  nearly 
a  right  angle.  This  gives  more  lift  to  the  outer  wing  than 
would  be  the  case  with  straight  entering  edges,  while  the 
inner  end  losses  increase  correspondingly  in  lift  and  hence 
tends  further  to  depress  the  inner  tip.  The  greatest  value 
of  sweep  back  is  found  in  its  resistance  to  side  slip.  If  the 
machine  should  be  ''over  banked,"  and  tend  to  slip  down 
toward  the  inner  side  of  the  turn,  the  backward  angle  of 
the  inner  wings  will  cause  the  entering  edge  to  meet 
the  side  stream  at  more  nearly  right  angles,  and  thus  tend 
to  reduce  the  bank  and  the  inner  slide  slip.  In  this  respect 
the  retreat  is  an  aid  to  lateral  stability.  On  the  other  hand, 
the  sweep  back  tends  to  keep  the  machine  rolling  in  rough 
weather  for  side  gusts  then  meet  the  inclined  wing  edges 
at  an  effective  angle.  This  is  particularly  noticeable  in 
landing. 

Raked  Tips.  This  subject  was  discussed  under  "Stand- 
ard Wing  Section,"  but  it  may  be  repeated  here,  that  Eiffel 
in  an  experiment  with  the  Coanda  Wing  (Eiffel  38),  found 
that  the  L/D  of  the  raked  wing  was  20  per  cent  higher 


PLAN  FORM 


187 


than  with  the  same  wing  in  rectangular  form.   This  value 
would  not  be  safe  to  assume  with  all  sections. 

Tandem  Arrangement.  In  tandem  wings  as  shown  by 
Fig.  9,  the  downward  wash  of  the  front  wing  (m)  will 
affect  the  rear  wing  (n)  by  causing  a  change  in  the  relative 


on  the  Rear  Wings  of  a  Tandem  Pai 


direction  of  the  flow.  If  the  front  wings  are  at  an  incidence 
of  (i)   degrees  (Fig.  11),  the  deviation  or  washdown  of 
the  air  stream  to  the  rear  will  be  expressed  by  :  d=  (0.5i+ 
1).     If  i'  =  incidence  of   rear   wing  measured   from   the 


Kig.   1 


-•     T^'if^ei?    Arrangements   Used    in    Eiffel   Experiments 
m  Straight  Line,  (2)  Rear  Wing  at  2.5%  (3)   Rear  Wi 


(1)    Chord; 
ing  at  5'. 


chord  of  the  front  wing,  then  the  angle  of  incidence  made 
by  the  rear  chord  to  the  horizontal  will  be :  I  =  (i-i') 
—  (0.5i-f  1),  where  I  is  the  incidence  of  the  rear  plane  with 
the  horizontal. 

Experiments  by  Eiffel  on  tandem  planes  with  circular 
cambered  aerofoils  gave  exceedingly  good  results  for 
certain  combinations  (Fig.  12).  These  arrangements  were 
used,  (1)  Chords  in  a  straight  line  (2),  Rear  aerofoil  tilted 
down  at  a  negative  angle  of  2.5°,  (3),  Rear  Plane  tilted 


188 


PLAN  FORM 


down  at  a  negative  angle  of  5°.  In  all  cases  the  camber 
was  1/13.5  of  the  chord,  and  the  front  and  rear  wings  were 
spaced  two  chord  widths  apart.  While  the  drag  did  not 
change  much  for  any  of  the  arrangements,  the  lifts  varied 
widely,  and  arrangement  (2)  is  by  far  the  more  efficient 
in  lifting  capacity.  No.  2  is  50  per  cent  greater  than 
(1),  and  has  twice  the  lifting  ability  of  (3).  For  the 
same  angle  of  incidence,  the  front  wing  does  the  same 
amount  of  lifting  in  all  cases,  the  difference  being  en- 
tirely due  to  the  changes  in  the  rear  surface.  In  (2) 
the  lift  of  the  rear  aerofoil  is  actually  13  per  cent  greater 
than  the  front  plane.  The  following  tables  give  the  results  : 


TABLE    OF    LIFTS    IN    GRAMS    FOR    THREE    ARRANGEMENTS. 


ANGLE  OF 
INCIDENCE 

ARRANGEMENT 
(1) 

ARRANGEMENT 

(2) 

ARRANGEMENT 
(3) 

3 
6 

9 

12 
Average  at  all 
angles     

665 

987 

1315 
1540 

1127 

0.64 

1094 
1568 
2068 
2326 

1764 
1.00 

334 
703 
965 
1347 

837 

Percentage  

0.47 

Fig.   13.     Drzewiecki  Tandem  Arrangement  for  Longitudinal  Stability. 


The  lifts  in  the  above  table  are  for  the  two  planes 
working  together,  and  the  angle  of  incidence  is  the  angle 
of  the  front  aerofoil,  or  rather  the  angle  of  the  combina- 
tion. The  wings  were  15x90  centimeters,  aspect  ratio  =  6. 


PLAN  FORM  189 

Fig.  12  shows  the  construction  clearly.   This  is  only  true 
for  circular  arched  surfaces  of  the  camber  given. 

M.  Drzewiecki  working  with  Eiffel's  results  on  the 
above  combinations,  produced  an  inherently  stable  tan- 
dem monoplane,  in  which  the  front  and  rear  wings  were 
of  different  cambers  and  were  set  at  different  incidences. 
The  front  wing  is  Eiffel  No.  8  set  normally  at  8°  inci- 
dence, and  the  rear  wing  is  Eiffel  No.  13-bis  (Bleriot 
11-bis),   set  normally  at   5°.     The   center  of  gravity  is 


Fig.  14.     Drzwiecki    Tandem    Wing    Arrangement    for    Stability. 

approximately  half-way  between  the  two  wings, 
and  the  front  is  smaller  than  the  trailing  surface.  Because 
of  the  difference  in  area,  the  lift  of  the  front  wing  varies 
less  rapidly  than  the  rear  when  the  angle  of  the  machine 
changes  because  of  disturbed  air.  Should  the  machine 
"head  up,"  the  rear  wing  increases  faster  in  lift  than  the 
front,  and  hence  restores  the  machine  to  a  horizontal 
position.  Should  the  front  surface  drop,  the  incidence  is 
reduced,  but  as  incidence  of  the  rear  wing  is  less  than  the 
front  (8°),  the  rear  wing  is  reduced  to  nearly  a  zero  angle 
of  incidence — (With  little  lift).  The  front  wing  is  still 
inclined  at  a  considerable  incidence:  (3°)  when  the  rear 
is  at  zero.  This  drops  the  rear,  and  raises  the  front  wing 
so  that  the  normal  attitude  is  restored.  Lateral  stability 
is  obtained  by  moving  the  two  halves  of  the  front  wing 
in  relation  to  one  another,  the  relative  movement  being 
similar  to  that  of  ailerons. 


190 


WING  CONSTRUCTION 


CHAPTER  IX. 
WING  CONSTRUCTION. 

General  Wing  Frame  Layout.  In  many  ways,  the  frame 
of  the  wing  is  one  of  the  most  important  structural  parts 
of  the  aeroplane.  It  not  only  maintains  the  proper  aero- 
dynamic form  of  the  aerofoil,  but  also  transmits  the  air 
pressure  and  lift  to  the  body  of  the  machine,  and  there- 
fore carries  the  entire  weight  of  the  aeroplane  when  in 
flight.  In  spite  of  the  heavy  loading  on  this  frame  it  has 
been  brought  to  a  remarkable  degree  of  strength  and 
lightness.  Not  only  is  ''Brute"  strength  necessary,  but 
it  must  also  be  rigid  enough  to  properly  retain  the  out- 
lines of  the  aerofoil  with  the  heaviest  loadings,  hence 
the  efficiency  of  the  aeroplane  greatly  depends  upon  the 
stiffness  as  well  as  strength.  The  contour  of  the  entering 
edge  must  be  particularly  accurate  and  well  supported 
since  it  is  at  this  point  that  the  greater  part  of  the  lift 
is  obtained,  and  where  a  slight  deviation  in  form  will 
materially  aft'ect  the  lift  and  drag. 

The  fabric  surface,  on  which  the  air  pressure  is  exerted, 
must  transmit  the  pressure  and  lift  to  the  main  struc- 
tural members  through  the  parts  that  give  form  to  the 
surface.  The  fabric  surfacing,  being  flexible  and  pliant, 
must  be  supported  at  frequent  intervals  by  the  forming 
members  which  in  effect  are  similar  to  the  joists  of  a 
floor  system.  The  forming  members  are  then  supported 
in  turn  by  longitudinal  beams,  or  girders,  that  transmit 
the  pressure  to  the  point  where  the  load  is  applied.  The 
girders  not  only  carry  the  lifting  force,  but  must  also 
take  care  of  the  drag  which  acts  at  right  angles  to  the 

191 


192  WING  CONSTRUCTION 

lift.  To  pass  girders  that  are  sufficiently  strong,  and  yet 
within  the  limits  of  weight,  through  the  narrow  space 
between  the  top  and  bottom  surfaces  of  the  wing  is  not 
always  the  simplest  of  problems. 

Figs.  1  and  2  show  typical  wing  frames  in  diagram- 
matic form,  the  upper  views  are  the  plans,  while  at  the 
bottom  are  sections  taken  through  the  wing.  The  outlines 
of  the  sections  are  curved  to  the  outlines  of  the  aerofoil 
adopted  for  the  wings,  and  after  this  outline  is  drawn 
out  to  scale,  we  must  maneuver  our  structural  members 
so  that  they  will  lie  entirely  between  the  surfaces. 

In  Fig.  1,  the  forming  ribs  are  indicated  by  R,  these 
being  the  members  curved  to  the  aerofoil  form.  They  are 
spaced  along  the  length  of  the  wing  at  intervals  of  about 
one  foot  and  the  fabric  is  applied  to  the  top  and  bottom 
edges  of  the  rib.  The  ribs  are  fastened  to  the  front  spar 
F,  and  the  rear  spar  S.  The  spars  are  equivalent  to  beams, 
and  are  for  the  purpose  of  transmittmg  the  lift  of  the 
ribs  to  the  body.  A  thin  strip  E  (nosing)  running  along 
the  entering  edge  of  the  wing,  serves  to  hold  the  fabric 
taut  at  this  point  and  also  forms  it  to  the  shape  of  the 
aerofoil  entering  edge.  The  thin  trailing  edge  strip  (T) 
performs  the  same  purpose,  and  the  wing  outline  is  com- 
pleted by  the  "End  bow"  (A)  which  retains  the  fabric 
at  the  wing  tips.  Between  the  front  spar  F  and  the  rear 
spar  S  is  the  trussed  "Drag  bracing,"  which  binds  the 
two  spars  into  a  truss  in  a  horizontal  direction,  and 
against  the  drag  of  the  surfaces.  This  consists  of  the 
"Drag"  wires  or  cables  (d)  and  the  short  wood  struts  (e), 
although  in  many  cases  the  ribs  are  strengthened  at  the 
point  of  attachment  of  the  drag  wires  and  serve  as  struts. 
The  aileron  G  is  located  at  the  outer  tip  and  is  hinged  to 
the  rear  spar  or  to  an  extension  of  the  rear  spar.  Between 
the  spars  are  thin  strips  known  as  "battens"  which  stiffen 
the  ribs  sideways,  these  are  shown  by  (F). 

Metal  connection  clips  C,  at  the  end  of  the  wing  spars, 


WING  CONSTRUCTION 


193 


are  for  attaching  the  wings  to  the  body,  or  for  connection 
of  the  two  halves  of  the  upper  wing  of  a  biplane.  Looking 
at  the  lower  sectional  view  we  see  the  interplane  struts 
of  a  biplane  attached  to  the  front  and  rear  spars  as  at  (m) 
and  (n).   Referring  to  the  plan  view,  the  location  of  the 


;     ■      'k  JBooy\ 


JFVGf.I. 


^^^EinEZngj^^ 


J^^.S. 


Fig.   1.     (Left)  Wing  Assembly  with  Spar  to  the  Rear  of  the  Entering  Edge. 
Fig.  2.  (Right)  Assembly  with  the  Front  Spar  at  the  Entering  Edge. 

Struts  is  indicated  by     *     *     *     at  the  points  where  the 
drag-bracing  is  attached  to  the  spars. 

Fig.  2  is  a  form  of  wing  in  which  the  spar  F'  also  forms 
the  entering  edge,  thus  eliminating  one  part  of  the  wing. 
One  objection,  to  this  construction  is  that  the  front  spar 
must  necessarily  be  shallower  than  the  spar  shown  in 


19-i 


WING  CONSTRUCTION 


Fig.  1.  The  rear  spar  is  in  the  usual  location  at  S',  the  two 
spars  being  connected  through  the  usual  end  bow  A'.  The 
trailing  edge  T'  may  be  either  a  thin  strip,  or  it  may  be  a 
thin  cable  as  indicated.  This  wing  is  similar  to  the  wing 
used  on  the  early  Wright  machines,  and  is  still  used  by 
Farman,  Voisin  and  other  European  manufacturers  of 
biplanes.  Usually  the  trailing  ends  of  the  ribs  overhang 
the  rear  spar  for  quite  a  distance,  in  this  type  of  wing,  giv- 
ing a  flexible  trailing  edge.  The  front  and  rear  interplane 


Fig.  3.     Sub-Rib  Construction,  the  Sub-Ribs   (r)   Are  Placed  at  the  Enter- 
ing Edge. 

Struts   (m)   and   (n)   are  shown,  the  former  connecting 
with  the  front  spar  at  a  point  near  the  entering  edge. 

Fig.  3  shows  the  usual  construction  except  that  short 
"Sub-ribs"  marked  (r)  are  placed  between  the  main  ribs 
R  at  the  entering  edge.  These  short  ribs  increase  the 
support  and  accuracy  of  the  curve  at  the  entering  edge, 
or  else  allow  wider  spacing  of  the  main  ribs  R.  The 
fabric  must  be  well  supported  at  this  point,  not  only  to 
maintain  the  best  efficiency  of  the  aerofoil,  but  to  relieve 
the  stress  on  the  fabric,  as  it  is  here  (Top  surface)  that 
the  greatest  suction  pressure  comes.  Should  there  be  a 
rip  or  tear  near  the  entering  edge,  in  the  lower  surface, 


WING  CONSTRUCTION 


195 


the  upper  fabric  will  be  subjected  to  both  the  pressure 
underneath  and  the  vacuum  above.  This  adds  fully  25 
per  cent  to  the  load  on  the  upper  facing. 

The  main  spars  may  be  of  wood  or  steel  tubing,  al- 
though the  former  material  is  generally  used.  They  are 
of  a  variety  of  forms,  the  *T"  beam  section,  solid  rec- 
tangular, hollow  box,  or  a  combination  of  plate  and  I  sec- 
tions, the  total  object  being  to  obtain  the  greatest  strength 
with  the  least  possible  weight.  When  made  up  of  sev- 
eral pieces  of  wood  they  are  known  as  "Built  up"  spars. 


^ 

CP                   c 

r^ 

p^-.^___^ 

L -n            1    t_^           1           —J =» 

4f         -^     4t 

Fig.  4.     Effects  of  CP.  Movement  on  Spar  Loading. 


The  load  on  the  spars  varies  with  the  total  weight  car- 
ried, and  also  with  the  movement  of  the  center  of  pres- 
sure due  to  changes  in  the  angle  of  incidence.  When 
the  center  of  pressure  moves  to  any  extent,  the  loads  on 
the  two  spars  may  vary  between  wide  limits,  and  in 
extreme  cases,  either  spar  may  carry  the  full  load.  This 
is  shown  clearly  by  Fig.  4,  a  section  taken  through  the 
wing.  The  front  spar  F  and  the  rear  spar  S  are  spaced 
by  the  distance  L,  the  respective  spar  loads  being  indi- 
cated by  Y  and  Z.  As  before  explained,  the  center  of 
pressure  moves  forward  at  large  angles   (CP),  while  at 


196  WING  CONSTRUCTION 

small  angles  it  moves  back  say  to  position  (CP').  Should 
it  move  back  as  far  as  CP-2,  the  load  will  come  directly 
under  the  rear  spar  and  this  member  will  therefore  carry 
the  entire  load.  When  at  the  forward  position  CP,  the 
greater  part  of  the  load  will  come  on  the  front  spar,  and 
only  a  small  portion  will  now  come  on  S.  In  the  same 
way,  when  at  a  small  angle  of  incidence,  the  center  of 
pressure  will  be  at  CP',  a  distance  (K)  from  the  rear 
spar.  The  greater  part  of  the  load  will  now  be  on  S. 
The  action  is  the  same  as  if  the  entire  weight  W  or  lift, 
were  concentrated  at  the  center  of  pressure. 

When  intermediate  between  the  two  spars,  the  center 
of  pressure  causes  a  bending  moment  in  the  rib  R,  and 
is  at  a  maximum  when  the  CP  is  midway  between  the 
two  spars.  It  will  be  seen  that  the  C.  P.  movement  has 
an  important  effect  outside  of  the  question  of  stability, 
and  this  travel  must  be  taken  into  careful  consideration 
when  the  strength  of  the  spars  is  calculated.  To  find  the 
load  on  the  rear  spar,  for  example,  with  the  center  of  pres- 
sure at  CP,  multiply  the  lift  W  by  the  distance  P,  and 
divide  by  the  spar  spacing  L.  This  will  give  the  load 
Z.  With  the  C.  P.  in  the  same  position,  the  load  on 
the  front  spar  will  be  the  difference  between  the  total 
lift  W  and  the  load  on  the  rear  spar,  or  Y  =  W  — Z.  With 
the  load  at  CP',  the  load  on  the  front  spar  will  be:  Y  = 
WxK,  and  the  load  on  S  will  be  Z  =  W  —  Y. 


L 

For  example,  we  will  assume  that  the  lift  W  =  1000 
pounds,  and  the  distance  P^  12  inches.  The  spar  spacing 
L  =  30  inches  and  the  center  pressure  is  at  CP.  The  load 
Z  on  the  rear  spar,  will  be : 
Z  =  WxP    =    1000x12  =  400    pounds.      The    load    on 


L  30 

the  front  spar  can  be  found  from  the  formula,  Y  =  W 
—  Z  =  1000  —  400  =  600  pounds. 


WING  CONSTRUCTION  197 

Fig.  5.  shows  a  typical  form  of  wing  construction  (rear 
spar  omitted).  The  front  spar  is  of  the  "Built  up  type," 
and  the  trailing  edge  is  a  flattened  steel  tube.  The  rear 
spar  is  simply  a  solid  rectangular  beam.  A  central  ash 
*T"  beam  is  used  as  the  front  spar,  with  vertical  spruce 
plates  on  either  side.  The  spruce  entering  edge,  or  "nos- 
ing," is  formed  to  the  shape  of  the  entering  edge  and  is 
hollowed  out  for  lightness.  The  rib  is  also  of  the  built 
up  type,  the  upper  and  lower  flanges  are  of  spruce  and 


Fig.   5.     Perspective  View  of  Wing  Construction  (Rear  Spar  Omitted),  Show- 
ing  Hollowed   Entering    Edge   and    Built-Up    Spar.      Rib    Is    of  the 
"I"   Beam   Type.      Courtesy    "Flight." 

the  middle  portion  (Web)  is  cotton-wood.  At  the  point 
where  the  spar  passes  through  the  rib,  the  rib  flanges 
pass  over,  and  are  tacked  to  the  spar.  The  spruce  nosing 
fits  closely  over  the  front  web  of  the  rib.  The  rib  flanges 
are  cut  away  so  that  the  outside  of  the  nosing  will  come 
flush  with  the  flange  line  of  the  rib. 

A  wing  of  decidedly  different  construction  is  the 
Caudron  monoplane  wing  shown  by  Fig.  6,  the  front  and 
rear  spars  of  this  wing  being  steel  tubes  with  an  entering 
edge  of  thin  wood.  The  drag  bracing  wires  may  be  seen 
connected  at  alternate  ribs  by  small  steel  plates  and  the 
latter  also  serve  to  attach  the  ribs  to  the  spars.  Instead  of 
being  cut  out  entirely,  the  webs  of  the  ribs  are  hollowed 


198  WING  CONSTRUCTION 

out  between  the  spars.  Probably  the  most  unique  feature 
is  the  construction  of  the  long  flexible  trailing  ends  of  the 
rib  at  the  right.  The  trailing  rib  edge  is  divided  into  an 
upper  and  lower  section  by  a  long  slot,  the  upper  sections 
being  rigid,  while  the  lower  edge  is  thin  and  flexible. 
The  flexible  edges  allow  the  lower  ends  of  the  ribs  to 
give  locally,  and  reduce  the  camber  when  struck  by  a 
heavy  gust.  This  aids  in  the  lateral  stability,  since  the 
lift  is  thus  considerably  reduced  at  the  point  of  impact,  and 


Fig.  6.  Caudron  Monoplane  Wing  with  Steel  Tube  Spars  and  Flexible 
Trailing  Edge.  A  Slot  in  the  Rear  of  the  Rib  Web  Permits  the 
Deflection  of  the  Trailing  Edge.  Drag  Wire  Bracing  Is  Used 
Between  the  Front  and  Rear  Spars.     Courtesy  "Flight." 

it  also  relieves  the  wing  of  unnecessary  stresses.  The 
rigid  upper  section  of  the  rib  acts  as  a  limit  stop  to 
the  lower  half,  and  prevents  the  flexure  from  exceeding 
a  certain  amount.  Owing  to  the  flexibility  of  the  trailing 
edges  a  steel  wire  or  cable  must  be  used  for  the  trailing 
edge. 

Details  of  the  framing  of  the  Standard  H-3  are  shown 
by  Fig.  7.  The  figure  at  the  left  gives  a  clear  idea  of 
the  connections  between  the  drag  struts  and  spar,  while 
the  view  at  the  right  shows  the  body  connection  at  the 


WING  CONSTRUCTION 


199 


end  of  the  spar.  I  am  indebted  to  "Aerial  Age"  for  these 
sketches.  The  main  spar  is  in  a  solid  piece,  channeled  out 
to  "I"  beam  form,  except  at  the  point  where  the  spruce 
drag  strut  is  attached.  At  the  end  of  this  strut  is  at- 
tached a  sheet  steel  fitting  that  affords  a  means  of  con- 
necting the  drag  wires,  and  for  fastening  the  strut  to  the 
main  wing  spar.  At  the  point  of  attachment,  wooden 
plates  are  fastened  to  either  side  of  the  spar.  These  pre- 
vent the  fitting  bolts  and  the  fitting  from  sliding  along 
the  spar  when  subjected  to  an  uneven  pull  in  the  wires. 
A  veneer  top  and  bottom  plate  still  further  strengthen 


Fig.  7.  Standard  H-3  Wing  Construction.  Spar,  Rib  and  Drag  Strut  Con- 
nections at  Left.  Body  Connection  Fitting  or  Hinge  at  Right.  Note 
Drag  Wire  Fittings.     Courtesy  "Aerial  Age." 


this  joint  and  hold  the  sub-rib  in  place.  The  main  ribs 
are  strengthened,  at  the  point  where  the  spar  passes 
through  the  rib  web,  by  small  vertical  blocks.  In  the 
right  hand  figure  the  steel  clevis  is  shown  bolted  to 
the  spar.  A  lug  for  the  wing  drag  wire  is  brought  out 
from  the  fitting.  The  clevis  on  the  wing  engages  with  a 
similar  clevis  on  the  body  of  the  machine,  and  the  two 
are  fastened  together  with  a  bolt  or  pin. 

Fig.  8  is  a  photograph  of  a  biplane  wing  with  the 
framing  members  completed  and  ready  for  the  applica- 
tion of  the  fabric.  At  the  right  is  the  opening  left  for 
the  aileron,  and  at  the  left  is  the  observation  port,  the 
latter  coming  next  to  the  body.   As  this  is  a  lower  wing, 


200 


WING  CONSTRUCTION 


the  sockets  for  the  attachment  of  the  interplane  struts  can 
be  seen  on  the  upper  and  near  edges  of  the  main  spar. 
Between  the  spars  are  very  thin  wood  strips  running 
with  the  length  of  the  wing.  These  are  the  ''Battens" 
used  for  stiffening  the  ribs  between  the  points  of  sup- 
port at  the  spars.  As  the  distance  between  the  spars 
is  comparatively  great,  in  respect  to  the  thickness  of 
the  rib  flanges,   some  sidewise  support  of  this  kind  is 


Fig.  8.  Typical  Biplane  Wing.  Gap  for  Aileron  Shown  at  Right  End  of 
Wing.  Left  End  Rests  Against  Fuselage,  the  Observation  Port 
Being  Cut  Out  at  the  Upper  Left  Ha«d  Corner.  Drag  Wire  Brac- 
ing Clearly  Shown.     Courtesy  "Aerial  Age." 

necessary.  The  drag-bracing  cables  cross  three  rib  spaces, 
or  are  fastened  to  every  fourth  rib.  Between  the  front 
spar  (at  the  bottom),  and  the  entering  edge,  are  the 
small  strips  that  serve  as  sub-ribs.  Double  cross  bracing 
is  used  at  the  inner  end  of  the  wing  (left),  while  addi- 
tional knee  braces  are  placed  at  the  aileron  opening, 
and  at  the  outer  tips.  This  is  necessary  to  withstand  the 
stresses  due  to  assembling  and  handling,  rather  than 
for  the  flight  stresses. 


WING  CONSTRUCTION 


201 


Fig.  9  is  a  Standard  Wing  ready  for  covering.  Before 
the  fabric  is  applied,  a  narrow  cloth  strip  is  wrapped 
over  the  trailing  edge,  as  shown,  and  is  stitched  to  the 
frame.  This  forms  a  means  of  stitching  the  main  covering 
at  the  rear  edge,  where  the  ends  of  the  upper  and  lower 
surfaces  meet. 

Wing  Fabric  or  Covering.  At  the  present  time  un- 
bleached Irish  linen  is  used  almost  exclusively  for  cover- 
ing the  wing  structure,  although  in  the  early  days  of 
flying  rubberized  fabrics  were  used  to  a  great  extent. 


Fig.  9.      Standard  H-3  Wing  Ready  for  Covering.     Opening  for  Aileron  Flap 
Shown  at  Upper  Left  Hand  Edge  (Trailing  Edge). 

After  the  linen  is  stretched  on  the  wing  frame,  it  is  given 
several  coats  of  a  special  preparation  commonly  known 
as  "Dope"  to  proof  the  fabric  against  moisture.  In  addi- 
tion to  waterproofing,  the  dope  adds  considerably  to  the 
strength  of  the  fabric  and  shrinks  it  tightly  on  the  ribs — 
much  more  evenly  than  could  be  done  by  hand.  When 
completely  "Doped,"  the  linen  should  be  proof  against 
the  effects  of  salt  water,  moisture,  or  extreme  dryness, 
and  the  fabric  must  be  "Drum  tight"  at  every  point 
on  the  surface  of  the  wing. 


202 


WING  CONSTRUCTION 


The  linen  should  have  a  tensile  strength  of  at  least 
75  pounds  per  inch  of  width  in  any  direction,  and  weigh 
from  3.75  to  4.4  ounces  per  square  yard.  It  must  be  wet 
spun,  free  from  filling  matter  and  uncalendered.  As  a 
usual  thing,  the  width  should  not  be  less  than  36  inches, 
although  the  width  can  be  altered  to  meet  conditions  of 
rib  spacing,  etc.  The  U.  S.  A.  seaplane  specifications 
(1916)  require  a  minimum  strength  along  the  warp  of 
75  pounds  per  inch  width,  and  85  pounds  per  inch  of 
width  along  the  weft.  The  following  table  gives  the 
properties  of  well  known  wing  fabrics : 


NAME  AND 

WEIGHT 
PER    SQ.    YD. 
IN   OUNCES 

g5 

THREADS    PER    INCH    WIDTH 

SPECIFICATION   NO. 

WARP 

WEFT 

U.  S.  A.— 1002 
Brit.  Gov.— A-37 
Brit.  Gov.— A-40 
McBratney 
Curtiss  Stand. 
R.  A.  F.— 17-C 

3.75-4.40 
4.12 
3.75 
4.00 
4.00 
4.00 

75 

112.0 
123.4 
85.0 
90.0 
92.0 

85 

115.6 
129.4 
108.0 
100.0 

95.0 

94 
95 

96 
92 

100 

99 

100 

95 

Wing  Dope.  AA^ing  dopes  are  in  nearly  every  case  based 
on  cellulose — either  cellulose  acetate  or  nitrate  being  the 
most  common  base.  This  has  proven  far  superior  to  the 
resin,  copal,  gum  or  oil  bases  contained  in  ordinary  var- 
nish, since  the  cellulose  of  the  dope  seems  to  amalgamate 
with  the  cellulose  of  the  flax  fiber  and  bond  the  whole 
into  an  integral  structure.  The  fact  that  the  dope  must 
be  elastic  bars  the  use  of  shellac  or  other  hard  resin 
solutions.  The  solvents  used  for  the  cellulose  dopes  vary 
with  the  makers,  some  using  amyl-acetate,  tetrach- 
lorethane,  etc.,  while  others  use  special  secret  compounds 
that  are  best  adapted  for  their  bases.  Many  of  the  solvents 
give  off  poisonous  gases  in  drying,  and  this  must  be 
guarded  against  by  good  ventilation.     The  vapor  of  te- 


WING  CONSTRUCTION 


203 


trachlorethane  is  particularly  dangerous,  and  has  resulted 
in  many  deaths. 

Doped  surfaces  have  from  10  to  25  per  cent  greater  ten- 
sile strength  and  resistance  to  tearing  than  the  undoped 
linen,  and  increases  the  weight  of  the  fabric  by  about  0.7 
ounce  per  square  yard  for  each  coat  applied.  Under  or- 
dinary weather  conditions,  dope  will  require  from  20  to  40 
minutes  per  coat  for  drying,  and  at  least  one-half  hour 


Fig.  10.  Complete  Framing  Plan  of  Typical  Monoplane  Structure.  CA) 
:d?^\'  ^^-^  F^xf^^",^^''-  (M)  Motor.  (S)  Stabilizer.  (E)  Elevator. 
(R)   Vertical  Rudder. 

should  be  allowed  between  each  coat.  Weather  condi- 
tions have  a  great  effect  on  the  action  of  dope,  and  with 
cellulose  compounds  the  best  results  are  obtained  in  a 
clean  dry  room,  well  warmed,  and  without  drafts.  On 
rainy  days  the  linen  is  damp  and  the  dope  does  not  set 
well,  and  this  trouble  is  not  greatly  helped  by  artificial 
heat.  Drafts  cause  white  spots  and  streaks,  especially 
if  cold  air  is  allowed  to  enter  directly  upon  the  warm  wing 
surface.    To  prevent  drafts  the  ventilating  ducts  should 


204  WING  CONSTRUCTION 

be  near  the  floor,  and  as  the  vapor  is  heavier  than  the 
air,  and  flows  downwards,  this  means  of  ventilation  is 
entirely  practicable. 

Applying  the  Dope.  The  number  of  coats  depend  upon 
the  character  of  the  job,  but  at  least  five  coats  should  be 
applied,  and  preferably  seven.  On  the  best  grade  of  work, 
the  dope  is  generally  covered  with  three  or  more  final 
coats  of  spar  varnish,  although  this  is  not  absolutely 
necessary.  For  ordinary  work,  dope  alone  on  Irish  linen 
has  proved  very  satisfactory  for  land  machines,  five  coats 
being  the  usual  amount  applied  on  exhibition  aeroplanes 
and  planes  for  amateur  use.  Seven  coats  of  dope  with 
three  coats  of  spar  varnish  are  specified  for  military  ma- 
chines that  are  to  be  used  on  salt  water.  Seaplanes  are 
subjected  to  conditions  that  are  particularly  hard  on 
fabric  and  must  be  protected  accordingly. 

In  applying  the  dope,  at  least  one-half  hour  should  be 
allowed  for  drying  between  each  coat,  and  more  if  pos- 
sible. The  first  two  coats  should  be  painted  on  lightly, 
the  purpose  being  simply  to  fill  the  pores  of  the  fabric 
and  to  prevent  the  succeeding  coats  from  sinking  through. 
If  the  first  two  coats  are  too  heavy,  the  dope  filters  through 
the  mesh  of  the  linen  and  drops  on  the  lower  surface, 
causing  spots  and  a  waste  of  a  very  expensive  material. 
Dope  is  expensive  even  with  the  greatest  care  exercised 
in  its  application,  and  the  writer  has  seen  cases  where 
the  first  two  coats  were  so  heavily  applied  that  fully  50 
per  cent  of  the  fluid  ran  through  and  caked  in  among  the 
structural  parts  of  the  machine.  This  ran  the  doping 
expense  up  to  a  terrific  figure.  The  cloth  should  be  dry, 
and  the  work  performed,  if  possible,  on  a  dry  day.  To 
save  dope,  never  take  out  of  the  supply  drum  more  than 
can  be  used  for  one  coat,  for  the  dope  soon  becomes  tacky 
on  exposure  to  the  air,  and  a  satisfactory  job  is  hard  to 
obtain  if  it  gets  in  this  condition. 

In  placing  the  fabric  on  the  wings,  particular  care  must 


WING  CONSTRUCTION 


!05 


be  taken  in  stretching  so  as  not  to  have  it  too  tight  when 
cellulose  dope  is  used.  The  dope  shrinks  the  linen  to  a 
very  considerable  extent,  and  if  it  is  too  tight  to  begin 
with,  the  stress  due  to  shrinkage  will  place  an  excessive 
stress  on  both  the  fabric  and  the  structure.  When  of  the 
proper  tautness,  the  fabric  should  sound  like  a  drum  when 
snapped  with  the  finger.  Any  less  tension  than  this  will 
permit  the  fabric  to  sag  badly  when  under  the  air  pres- 
sure and  reduce  the  efficiency  of  the  wings.    In  fastening 


Fig.  10a.  Testing  the  Wing  Structure  of  a  Curtiss  Biplane  by  Means  of 
Sand  Bag  Loading.  The  Wings  Are  Turned  Upside  Down  and  a 
Sand  Load  Is  Laid  Uniformly  Over  the  Wings  So  That  They 
Produce  a  Load  Equal  to,  or  Greater  Than,  the  Flight  Load. 


the  cloth  it  should  be  just  stretched  taut  and  no  more. 
In  damp  weather  the  cloth  can  be  stretched  a  little  tighter 
than  in  dry  weather. 

Transparent  Coverings.  In  some  types  of  battle-planes 
and  scouts,  a  part  of  the  wing  section  directly  above  the 
pilot  is  covered  with  a  transparent  fireproof  cellulose 
sheet,  much  resembling  celluloid.  This  permits  the  pilot 
to  see  above  him  through  the  overhanging  wing,  and  is 


206  WING  CONSTRUCTION 

of  great  value  in  action.  In  some  cases,  a  strip  is  placed 
on  the  lower  wings  along  the  sides  of  the  body  so  that 
the  ground  is  also  easily  visible.  These  cellulose  sheets 
will  not  crack  nor  splinter,  and  are  nearly  as  flexible  as 
rubber.  Celluloid  of  the  ordinary  variety  must  not  be 
used,  for  this  is  easily  ignited  and  is  likely  to  start  a 
disastrous  fire. 

Placing  the  Fabric  on  the  Wings.  In  some  aeroplanes 
the  seams  of  the  fabric  are  run  parallel  to  the  ribs,  and  are 
tacked  or  sewed  directly  to  them,  while  in  other  cases  the 
seams  are  run  diagonally  across  the  plane  or  on  the  ''bias." 
Diagonal  seams  are  most  satisfactory,  and  if  care  is  taken 
there  is  no  more  waste  of  linen  than  with  the  straight  seam. 
The  seams  should  be  of  the  double-lapped  or  "English 
welt"  type,  and  this  of  course  necessitates  sewing  before 
the  fabric  is  placed  on  the  wings.  The  seams  used  on  over- 
coats are  satisfactory  for  this  purpose,  and  give  a  cover- 
ing that  wall  not  stretch  nor  bag.  Some  use  linen  thread 
and  others  use  silk,  but  the  linen  is  preferable,  since  dope 
often  causes  silk  to  rot.  The  seams  should  be  covered 
with  linen  to  protect  them  from  the  weather  and  to  pre- 
vent the  entrance  of  water  to  the  interior  woodwork. 

Ordinarily,  the  wing  is  turned  upside  down  for  cover- 
ing, with  the  concave  side  uppermost.  The  seams  are 
sewed  together  so  that  the  completed  fabric  is  wider  than 
the  length  of  the  wing  and  is  a  little  longer  than  is  neces- 
sary to  wrap  entirely  around  the  width  of  the  wing.  The 
fabric  is  then  temporarily  fastened  along  the  trailing  edge, 
is  passed  under  the  wing  to  the  front  edge,  and  over  the 
concave  upper  side  back  to  the  trailing  edge.  At  this 
point  the  excess  material  will  hang  down  over  the  rear 
edge.  With  the  wing  in  its  upside  down  position,  the  con- 
vex side  will  be  at  the  bottom,  and  if  a  weight  is  hung 
on  the  overhanging  material  at  the  rear  edge,  the  cloth 
will  be  pulled  tight  against  the  low^er  convex  side  and 
straight  across  and  above  the  concave  side.     The  fabric 


WING  CONSTRUCTION 


207 


at  the  top  is  then  stretched  along-  the  cordal  line  of  the 
ribs.  By  laying  a  narrow  board  on  top  of  the  fabric,  and 
near  the  entering  edge,  the  fabric  can  be  brought  down 
uniformly  along  the  concave  edge  of  the  ribs,  and  by  tack- 
ing or  sewing  as  the  board  is  moved  back  the  concave 
face  can  be  covered  without  further  trouble.  After  the 
concave  face  is  disposed  of,  the  wing  can  be  turned  over 
and  the  fabric  is  then  fastened  to  the  convex  side  of  the 
ribs. 

One  method  of  fastening  the  linen  is  to  lay  tape  over 


Fig.  n.  Method  of  Stretching  Fabric  on  Wings.  Fabric  Passes  Under  and 
Then  Over  Concave  Side  and  Is  Pressed  Down  into  Hollow  by  a 
Board  as  Shown. 


the  ribs,  and  then  drive  tacks  through  the  tape  and  fabric 
into  the  rib.  The  tape  keeps  the  tacks  from  tearing 
through  the  linen.  The  tape  should  be  heavy  linen  of 
from  ^  to  1^  inches  wide,  and  laid  in  cellulose  before 
tacking,  so  that  the  tape  will  be  cemented  to  the  fabric 
and  the  solution  will  be  driven  into  the  tack  holes.  After 
the  tape  is  in  place,  it  should  be  covered  with  not  less  than 
three  coats  of  cellulose  dope  before  the  main  surface  is 
treated.  This  gives  an  additional  three  coats  over  the 
tape  where  it  is  most  needed  for  protection  against  mois- 
ture.   In  any  case,  the  seam  or  tacking  strip  should  be 


208  WING  CONSTRUCTION  ' 

pressed  down  so  that  it  projects  as  little  as  possible  above 
the  general  surface  of  the  wing. 

Tacking  is  not  desirable,  for  tacks  and  nails  always 
tend  to  split  the  thin  members  of  the  rib,  and  very  often 
corrode  the  cloth  and  weaken  the  fabric.  This  has  re- 
sulted in  the  whole  fabric  being  ripped  off  while  in  flight. 
Iron  or  steel  tacks  should  never  be  used,  as  they  destroy 
the  strength  of  the  fabric  very  rapidly  by  the  formation 
of  rust,  and  particularly  with  sea-planes  used  on  salt 
water.  Sewing  the  fabric  to  the  ribs  with  linen  thread 
is  the  most  satisfactory  method  and  is  in  general  use 
among  high-grade  builders.  if 

The  fabric  can  be  stitched  to  bands  of  thread  or  tape, 
the  latter  being  wrapped  about  the  rib.  Stitches  can  also 
be  taken  from  one  surface  straight  through  to  the  other. 
The  thread  or  tape  bands  on  the  ribs  are  merely  wrappings 
taken  around  the  rib  flange,  and  through  the  lightening 
holes,  these  bands  usuallybeing  about  4  inches  apart.  They 
at  once  tie  the  flange  to  the  web  and  form  a  soft  surface 
into  which  to  take  the  stitches.  Fig.  4  in  Chapter  X 
shows  the  thread  bands  (d)  wrapped  around  the  rib  flange 
(G),  and  through  the  lightening  hole,  the  fabric  lying 
above  and  below  the  rib  flanges  as  shown.  A  section 
view  through  the  rib  and  fabric  is  shown  at  the  right. 

Varnishing.  When  varnish  is  to  be  used  over  the  dope, 
only  the  best  grade  of  spar  varnish  should  be  used,  since 
any  other  kind  is  soon  destroyed  by  moisture.  From  two 
to  three  varnish  coats  will  be  sufficient,  and  each  coat 
should  be  thoroughly  dried  and  sandpapered  before  the 
next  is  applied,  care  being  taken  not  to  injure  the  fabric 
with  the  sandpaper.  Sandpapering  between  the  dope  lay- 
ers is  not  necessary,  since  each  successive  coat  partially 
dissolves  the  preceding  coat,  and  thus  welds  the  layers 
together.  Varnish,  however,  does  not  act  in  this  way, 
and  the  coats  must  be  roughened.  Shellac  rots  the  linen 
and  should  not  be  used. 


WING  CONSTRUCTION  209 

The  Government  specifies  one  coat  of  flexible  white 
enamel  in  which  a  small  quantity  of  lead  chromate  is 
mixed.  This  is  applied  over  the  last  coat  of  varnish.  The 
lead  chromate  filters  out  the  actinic  rays  of  the  sun,  thus 
reducing  the  injurious  efifect  of  the  sunlight  on  the  cover- 
ing.   If  coloring  matter  is  added  to  the  dope,  it  should  be 


Fig.   12.     Wing  Structure  of  Handley-Page  Giant  Biplane.     Courtesy  "Aerial 
Age." 

in  liquid  form,  as  powders  destroy  the  strength  and  tex- 
ture of  the  dope  deposit. 

Patching  Fabric.  The  majority  of  dopes  can  be  used 
as  cement  for  patching,  but  as  dope  will  not  stick  to  var- 
nish, all  of  the  varnish  around  the  patch  must  be  thor- 
oughly removed  with  some  good  varnish  remover.  The 
varnish  must  be  thoroughly  cleaned  ofif  or  there  will  be 
no  results.  Before  applying  the  dope,  the  patch  must  be 
well  stitched  all  around  the  edges,  then  cemented  with  the 
dope.  The  patch  must  now  be  covered  with  at  least  five 
coats  of  thin  dope  as  in  finishing  the  surface.  Particular 
attention  must  be  paid  to  filling  the  dope  in  the  stitching. 


CHAPTER  X. 
WING  CONSTRUCTION  DETAILS. 

Types  of  Ribs.  The  rib  first  used  by  the  Wright  Broth- 
ers consisted  of  two  spruce  strips  separated  by  a  series 
of  small  pine  blocks.  Practically  the  same  construction 
was  used  by  Etrich  in  Austria.  With  the  coming  of  the 
monoplane,  and  its  deep  heavy  spars,  the  old  Wright  rib 
was  no  longer  suitable  for  the  blocks  were  not  efficient 
in  thick  wing  sections.  The  changes  in  the  wing  form 
then  led  to  the  almost  universal  adoption  of  the  "V 
type  rib  in  which  an  upper  and  lower  flange  strip  are 
separated  by  a  thin  vertical  web  of  wood.  At  present 
the  "V  rib  is  used  on  nearly  every  well  known  machine. 
It  is  very  strong  and  light,  and  is  capable  of  taking  up 
the  end  thrust  of  the  drag  wires,  as  well  as  taking  care  of 
the  bending  stresses  due  to  the  vertical  loading  or  lift. 

Fig.  1  shows  the  original  Wright  rib  with  the  "Battens" 
or  flanges  (g)  and  the  spacing  blocks  B.  The  front  spar 
is  at  the  leading  edge  (F),  and  the  rear  spar  at  S.  An  *T"^ 
beam,  or  "Monoplane"  type  is  shown  by  Fig.  2,  and  as 
will  be  seen  is  more  suitable  for  deep  spars  such  as  (F') 
and  S'.  The  upper  and  lower  flanges  (g)  are  separated 
by  the  thin  perforated  web  (w),  the  sectional  view  at  the 
right  showing  the  connection  between  the  flanges  and 
the  web.  Lightening  holes  (h)  reduce  the  weight  of  the 
web,  with  enough  material  left  along  the  center  of  the 
web  to  resist  the  horizontal  forces.  The  web  is  glued  into 
a  slot  cut  in  the  flanges,  and  the  flanges  are  then  either 
tacked  to  the  web  with  fine  nails,  or  bound  to  it  by  turns 
of  thread  around  the  flange. 

210 


WING  DETAILS 


211 


On  the  average  machine,  the  web  is  about  3/16  thick, 
while  the  flanges  are  from  3/4  to  1  inch  wide,  and  from 
3/16  to  1/4  inch  thick.  On  the  very  large  machines,  the 
dimensions  of  course  are  materially  increased.  At  the 
strut  locations  in  biplanes,  and  the  point  of  cable  bracino- 
attachment  in  monoplanes,  the  ribs  are  increased  in 
strength  unless  the  end  thrust  of  the  stay  wires  is  taken 
up  by  a  separate  strut.  At  the  point  of  stay  connection 
in  the  old  Nieuport  monoplane  the  rib  was  provided  with 


z^-? 


r- 


\ 


rf'     r9' 


r- 


Ck 


t:^ 


TBe^i^  T^yJF^ 


Fig.    1.     ^yright  Type  Rib  with  Battens  and  Block  Separators.     Fig   2      Mono- 
plane or  "I"  Type  of  Rib  with  Solid  Web. 

a  double  web  thus  making  a  hollow  box  form  of  section 
great  enough  to  account  for  the  diagonal  pull  of  the 
stays. 

Fig.  3  shows  the  Nieuport  monoplane  ribs,  which  are 
good  examples  of  box  ribs.  The  sections  at  the  left  are 
taken  through  the  center  of  the  ribs.  The  wing  chord 
tapers  from  the  body  to  the  wing  tip,  while  the  thickness 
of  the  wing  section  is  greatest  at  the  middle,  and  tapers 
down  both  toward  the  tips  and  toward  the  body.  The 
upper  section  is  located  at  the  body,  the  second  is  located 
midway  between  the  body  and  the  tip,  and  the  other  two 


212 


WING  DETAILS 


are  near  the  tips,  the  bottom  being  the  last  rib  at  the 
outer  end.  The  ribs  shown  are  of  box  form  as  they  are 
at  points  of  connection,  but  the  intermediate  ribs  are  of 
the  'T"  type  shown  by  Fig.  2. 

Rib  Material.  In  American  aeroplanes,  the  flanges  of 
the  ribs  are  generally  made  of  spruce.  The  webs  are  of 
poplar,  whitewood,  cottonwood  or  similar  light  material. 
There  is  not  a  great  deal  of  stress  on  a  rib,  and  the  strong- 
est material  is  not  necessary,  but  as  there  are  a  great 


Nieuport  Monoplane  Ribs.  This  Wing  Is  Thickest  in  the  Center 
and  Washes  Out  Toward  Either  End,  Thus  Making  All  of  the  Ribs 
Different  in  Curvature  and  Thickness.  At  the  Point  of  Stay  Wire 
Attachment  Double  Webbed  Box  Ribs  Are  Used. 

many  ribs  in  a  wing  assembly  lightness  is  a  primary 
consideration.  A  few  ounces  difference  on  each  rib  makes 
a  great  deal  of  difference  in  the  total  weight,  especially 
when  there  are  80  or  more  ribs  in  a  complete  machine. 
Exception  to  the  above  materials  will  be  found  in  the 
Curtiss  **Super-American"  Flying  Cruiser  which  has  ribs 
with  pine  webs  and  birch  flanges.  European  aeroplane 
practice  makes  use  of  hardwood  in  the  ribs. 

Web  Stififeners.    The  webs  being  thin  and  deep,  and 
cut  for  lightening  as  well,  need  bracing  at  the  points 


WING  DETAILS 


213 


where  concentrated  loads  are  placed,  such  as  at  the  front 
and  rear  spars,  and  at  points  between  the  lightening  holes. 
By  gluing  thin  strips  to  the  webs  (in  a  vertical  direc- 
tion), and  so  that  the  tops  and  bottoms  of  the  strips  come 
tight  against  the  upper  and  lower  flanges,  a  great  deal 
of  the  strain  on  the  web  can  be  avoided.  The  stiffening 
blocks  are  shown  by  (x)  in  Fig.  4,  and  are  placed  on  both 
sides  of  the  front  and  rear  spars  F  and  S,  and  also  be- 
tween the  lightening  holes  H. 

Flange  Fastenings.    In  the  section  at  the  right  of  Fig.  4 


Fig.  4.     Details  of  "I"  Type  Rib,  Showing  Lightening  Holes  and  Stiffeners., 

it  will  be  seen  that  the  web  is  inserted  into  a  groove  cut 
in  the  flanges  and  is  then  glued  into  place.  It  would  be 
unsafe  to  trust  entirely  to  the  glue,  owing  to  the  effects  of 
aging,  moisture  and  heat,  and  consequently  some  addi-^ 
tional  means  of  fastening  must  be  had.  It  has  been  cus- 
tomary to  nail  through  the  flange  into  the  web,  but  as 
the  web  is  only  about  3/16  inch  thick  it  is  likely  to  split. 
An  approved  method  is  shown  in  Fig.  4  in  which  Irish 
linen  thread  wraps  (d)  are  passed  through  the  lightening 
holes  (H),  and  over  the  flanges  (G).  The  thread  is  coated 
with  glue  before  wrapping  and  after,  and  when  dry  it 


214  WING  DETAILS 

is  thoroughly  varnished  for  protection  against  moisture. 
The  bands  are  spaced  from  3  to  4  inches  apart.  If  nails 
are  used  they  should  be  brass  nails — never  steel  or  iron. 

At  the  points  (F)  and  (S)  where  the  spars  pass  through 
the  web,  the  web  is  entirely  cut  out  so  that  the  flanges 
ordinarily  lie  directly  on  the  spars.  In  this  case  it  is 
necessary  to  bevel  the  spar  so  that  it  at  least  approxi- 
mately fits  the  curve  of  the  flange.  Sometimes  when  a 
full  size  spar  is  impossible,  as  in  cases  where  the  spar 
tapers  toward  the  tips,  wood  packing  pieces  may  be 
placed  between  the  flange  and  the  spar ;  tapered  to  make 
up  for  the  curve.  The  flanges  in  any  case  must  be  se- 
curely fastened  to  the  spar  by  brass  wood  screws  as  at 
(e),  and  the  edges  of  the  w^eb  should  fit  tightly  against 
the  sides  of  the  spars. 

Wing  Battens.  The  wing  battens  run  along  the  length 
of  the  wings,  from  end  to  end,  and  between  the  spars, 
and  serve  to  brace  the  ribs  sideways  as  shown  in  some 
of  the  general  views  of  the  wing  assembly.  To  accommo- 
date the  battens,  the  openings  (f)  are  cut  directly  under 
the  flange.  Usually  the  battens  are  thin  spruce  strips 
from  3/16  to  1/4  inch  thick  and  1/2  inch  wide,  and  should 
be  run  through  the  web  at  a  point  near  the  stiffeners. 
The  thickness  from  the  top  of  the  flange  to  the  under 
side  of  the  lightening  hole  is  from  1/2  to  3/4  inch  as 
indicated  by  (C). 

Strength  of  Wood  Ribs.  The  strength  of  a  rib  for  any 
individual  case  can  be  found  by  the  method  used  in  com- 
puting beams,  the  rib  usually  being  assumed  to  have 
a  uniformly  distributed  load,  although  this  is  not  actu- 
ally the  case,  as  before  explained.  The  greater  part  of 
the  load  in  normal  flight  is  near  the  front  spar,  but  this 
shifts  back  and  forth  with  the  angle  of  incidence  so  that 
there  is  no  real  stationary  point  of  application,  and  the 
rib  must  be  figured  for  the  maximum  condition.  The 
total  load  carried  by  one  of  the  intermediate  ribs  is  due 


WING  DETAILS  215 

to  the  area  between  the  ribs,  or  to  the  unit  loading  multi- 
plied by  the  rib  spacing  and  chord.  The  portion  of  the 
rib  between  the  spars  can  be  calculated  as  a  uniformly 
loaded  beam,  supported  at  both  ends.  The  entering  edge 
in  front  of  the  spar,  and  the  trailing  edge  to  the  rear 
may  be  taken  as  uniformly  loaded  beams  supported  at 
one  end.  The  proportion  of  the  loads  coming  on  the  ends 
and  center  position  can  be  taken  from  the  pressure  dis- 
tribution diagrams  as  shown  under  "Aerofoils." 

A  number  of  tests  were  made  on  ribs  by  Mr.  Heinrich 
of  the  Heinrich  Aeroplane  Company,  the  ribs  being  built 
up  on  short  pieces  of  spar  so  that  actual  conditions  were 
approached.  Instead  of  using  a  distributed  load,  such  as 
usually  comes  on  the  rib,  a  concentrated  load  was  placed 
at  the  center.  If  the  rib  were  uniform  in  section  the 
equivalent  uniformly  distributed  load  could  be  taken  as 
one-half  the  concentrated  load,  but  because  of  the  light- 
ening holes  this  would  not  be  very  exact.  It  would  be 
on  the  safe  side,  however,  as  such  a  test  would  be  more 
severe  than  with  a  uniform  load.  The  ribs  were  of  the 
same  type  as  shown  in  Fig.  5,  and  were  placed  32.5  inches 
apart.  The  front  spar  was  2  7/16  inches  deep,  the  rear 
spar  2  1/16  inches  deep,  and  the  overall  depth  of  the  rib 
at  the  center  was  3  1/8  inches.  The  rib  flanges  were  of 
white  wood  3/4  inch  wide,  and  3/16  inch  thick.  In  rib 
No.  1  the  web  was  solid  whitewood,  3/16  inch  thick,  and 
in  ribs  Nos.  2  and  3,  the  webs  were  mahogany  three-ply 
veneer  (5/32  inch  thick). 

Test  of  Rib  No.  1.  There  are  5  lightening  holes  between 
spars,  with  2  inches  of  material  left  between  the  holes, 
and  1/2  inch  between  first  hole  and  front  spar  opening. 
With  95  pounds  concentrated  load  at  the  center,  the  first 
fupture  appeared  as  a  split  between  the  first  hole  and 
spar  opening.  At  119  pounds,  the  flanges  had  pulled  away 
from  one  side  of  the  spar,  and  1/8  inch  away  from  the  web. 
Full   failure  at   127.5   pounds.    Web  was  split  between 


216  WING  DETAILS 

each  lightening  hole  with  a  complete  cross  break  at  center 
of  web,  the  latter  being  caused  by  a  brad  hole  in  the  web. 

Test  of  Rib  No.  2.  Laminated  web,  with  no  brads 
driven  opposite  lightening  holes.  At  140  pounds  rib  de- 
flected 5/16  inch,  and  when  relieved  sprang  back  only 
3/16  inch.  With  175  pounds,  the  deflection  again  was 
5/16  inch,  but  the  rib  continued  to  bend  slowly,  the  flanges 
pulling  away  from  the  web  and  spars.  The  wood  was 
not  broken  anywhere,  the  failure  being  in  the  brads  and 
glue. 

Test  of  Rib  No.  3.  Same  materials  as  No.  2,  but  web 
was  fitted  inside  of  the  'T"  beam  spar  and  the  rib  flanges 
were  screwed  to  the  spar.  At  175  pounds  there  was  no 
sign  of  rupture  anywhere,  and  the  deflection  was  5/16  inch. 
At  185  pounds  the  rib  broke  very  suddenly  and  cleanly, 
and  in  such  a  w^ay  as  to  indicate  that  this  was  the  true 
strength  of  the  rib.  The  normal  loading  on  the  rib  in  flight 
was  17.5  pounds,  uniformly  distributed,  so  that  with  a 
concentrated  load  of  370  pounds  equivalent,  the  safety 
factor  was  21.1. 

The  conclusions  to  be  arrived  at  from  this  test  are  as 
follows : 

(1)  When  a  solid  soft  wood  web  is  used,  there  should 
be  at  least  2  1/2  to  3  inches  between  lightening  holes. 

(2)  A  laminated  or  three-ply  web  is  the  best. 

(3)  No  brads  should  be  driven  opposite  lightening 
holes. 

(4)  The  web  should  fit  closely  to  the  spar  sides  and 
the  flange  of  the  rib  should  be  tightly  screwed  to  the  top 
and  bottom  of  spar. 

The  above  gives  an  idea  as  to  the  strength  of  the  usual 
form  of  wood  rib,  and  can  be  used  comparatively  for  other 
cases  if  the  reader  is  not  familiar  with  strength  calcula- 
tions. 

Making  the  Rib.  Wooden  webs  are  cut  out  on  the  band 
saw,  and  the  webs  are  so  simple  that  there  is  not  much 


WING  DETAILS 


217 


more  to  be  said  on  the  subject.  The  flanges,  however, 
must  be  steamed  and  bent  to  the  nearly  correct  form  be- 
fore assembly.  After  planing  to  size  and  cutting  the 
groove  for  the  reception  of  the  web,  the  ribs  are  placed 
in  the  steamer  and  thoroughly  steamed  for  at  least  an 
hour.  A  rib  flange  press  shown  by  Fig.  7  consists  of  two 
heavy  blocks  with  the  inner  faces  cut  approximately  to  the 
rib  outline.  The  steamed  ribs  are  then  placed  between 
the  blocks,  the  bolts  are  screwed  down  tight,  and  is  left 


Fig.   7.     Rib  Bending  Press   for  Curving  the  Rib  Flanges. 


for  24  hours  so  that  the  strips  have  ample  time  to  cool  and 
dry.  For  the  amateur  or  small  builder,  the  steamer  can 
be  made  of  a  galvanized  ''down-spout"  connected  with 
an  opening  cut  in  the  top  of  an  ordinary  wash  boiler. 
One  end  of  the  spout  is  permanently  sealed,  while  the 
other  is  provided  with  a  removable  cover  so  that  the 
strips  can  be  inserted.  A  hole  cut  near  the  center  of 
the  spout  is  connected  to  the  opening  in  the  boiler  cover 
by  a  short  length  of  spouting  or  pipe.  The  spout  should 
be  made  large  enough  in  diameter  to  contain  all  of  the 
ribs  that  can  be  pressed  at  one  time,  and  should  be  long 


218  WING  DETAILS 

enough  to  accommodate  longer  pieces  such  as  the  fuselage 
longerons,  etc. 

When  removed  from  the  press,  the  rib  flanges  can  be 
glued  to  the  webs  taking  care  that  the  glue  is  hot,  and 
that  it  thoroughly  covers  the  groove  surface.  The  rib 
must  now  be  held  accurately  in  place  in  a  second  form,  so 
that  the  true  rib  outline  will  be  retained  until  the  glue 
drys.  A  great  deal  depends  upon  the  accuracy  of  the 
second  form,  and  the  accuracy  with  which  the  web  outline 
is  cut.  The  larger  manufacturers  use  metal  rib  forms  or 
''jigs,"  but  the  small  builder  must  be  content  with  a 
wooden  form  consisting  of  a  board  fitted  with  suitable 
retaining  cleats,  or  lugs.  The  outline  of  the  aerofoil  is 
drawn  on  the  board,  the  tips  of  the  cleats  are  brought 
to  the  line,  and  are  screwed  to  the  board  so  that  they 
can  be  turned  back  and  forth  for  the  admission  and  re- 
lease of  the  ribs.  The  strip  bending  press  in  Fig.  7  is 
only  intended  to  bend  the  flanges  approximately  to  form, 
and  hence  two  layers  may  be  put  in  the  press  at  one 
time  without  much  error. 

Wing  Spars.  In  American  aeroplanes  these  members  are 
usually  of  the  solid  'T"  form  for  medium  size  exhibition 
and  training  machines,  but  for  small  fast  aeroplanes, 
where  every  ounce  must  be  saved,  they  are  generally  of 
the  built  up  type,  that  is,  made  up  of  two  or  more  mem- 
bers. In  Europe,  built  up  construction  is  more  common 
than  in  this  country,  and  is  far  preferable  for  any  ma- 
chine that  justifies  the  additional  time  and  expense. 
The  wing  spars  are  the  heaviest  and  most  important 
members  in  the  wing  and  no  trouble  should  be  spared 
to  have  them  as  light  as  the  strength  and  expense  will 
permit.  They  are  subjected  to  a  rather  severe  and  com- 
plex series  of  stresses ;  bending  due  to  the  load  carried 
between  supports,  compression  due  to  the  pull  of  the 
stay  wires,  bending  due  to  the  twist  of  non-central  wire 
fittings,  stresses  due  to  drag  and  those  caused  by  sudden 


WING  DETAILS 


219 


deviations  in  the  flight  path  and  by  the  torque  of  the 
motors.  These  should  be  accurately  worked  out  by  means 
of  stress  diagrams  if  the  best  weight  efficiency  is  to  be 
obtained. 

A  number  of  different  wing  spar  sections  are  shown 
by  Figs.  9,  10,  11.  Spar  (A)  in  Fig.  9  is  the  solid  one 
piece  *T"  type  (generally  spruce),  channeled  out  along 
the  sides  to  remove  the  inefficient  material  at  the  center. 
The  load  in  this  case  is  assumed  to  be  in  a  vertical  direc- 
tion. In  resisting  bending  stresses,  it  should  be  noted 
that  the  central  portion  of  the  material  is  not  nearly  as 
effective  as  that  at  the  top  and  bottom,  and  that  the  same 


C  B  A 

Fig.  9.     Types  of  Wing  Spars.     (A)  Is  the  "I"  Beam  Type.     CB)  Box  Spar. 
(C)  Is  Composite  Wood  and  Steel,  Wrapped  with  Tape. 

weight  of  material  located  top  and  bottom  will  produce 
many  times  the  results  obtained  with  material  located 
along  the  center  line.  At  points  of  connection,  or  where 
bolts  pass  through  the  spar,  the  channeling  is  discon- 
tinued to  compensate  for  the  material  cut  away  by  thb 
bolt  and  fittings. 

Spar  (B)  is  of  the  hollow  type,  made  in  two  halves 
and  glued  together  with  hardwood  dowel  strips.  The 
doweling  strips  may  be  at  the  top  and  bottom  as  shown, 
or  on  the  horizontal  center  line  as  shown  by  Spar  (J). 
The  material  of  the  box  portion  is  generally  of  spruce. 
This  is  a  very  efficient  section  as  the  material  lies  near 
the  outer  edge  in  every  direction,  and  offers  a  high  re- 


220 


WING  DETAILS 


sistance  to  bending,  both  horizontally  and  vertically.  Un- 
fortunately a  great  deal  depends  upon  the  glued  joints, 
and  these  require  careful  protection  against  moisture. 
There  is  absolutely  no  means  of  nailing  or  keying  against 
a  slipping  tendency  or  horizontal  shear.  The  best  ar- 
rangement to  insure  against  slipping  of  the  two  halves  is 
to  tape  around  the  outside  as  shown  by  spar  (E).  This 
is  strong  linen  tape  and  is  glued  carefully  to  the  spar, 
and  the  whole  construction  is  proofed  against  moisture 
by  several  coats  of  spar  varnish  and  shellac.  In  addition 
to  the  strengthening  effect  of  the  tape,  it  also  prevents  the 
wood  from  splintering  in  accidents. 


G  FED 

Fig.   10.     Four  Types  of  Wing  Spars,  the  Spar  D  Being  a  Simple  Steel  Tube 
as  Used  in  the  Caudron  and  Breguet  Machines. 


Spar  (C)  consists  of  a  central  ash  "I"  section,  with  steel 
strips  in  the  grooves.  Two  spruce  side  strips  are  placed 
at  either  side  as  stiffeners  against  lateral  flexure,  and  the 
•entire  construction  is  taped  and  glued.  This  is  very  ef- 
fective against  downward  stresses,  and  for  its  strength 
is  very  compact.  Since  spruce  is  much  stififer  than  either 
the  thin  steel  strip,  or  the  ash,  it  is  placed  on  the  outside. 
Spar  (D)  in  Fig.  10  has  been  described  before. 

Fig.  (F)  consists  of  two  spruce  channels  placed 
back  to  back,  with  a  vertical  steel  strip  between  them. 
Again  the  spruce  is  used  as  a  side  stiffener,  and  in  this 
case  probably  also  takes  a  considerable  portion  of  the 
•compression  load.  Spar  (G)  is  a  special  form  of  box  spar 
dsed  when  the  spar  is  at  the  entering  edge  of  the  wing. 


WING  DETAILS 


221 


the  curved  nose  being  curved  to  the  shape  of  the  aerofoil 
nose.  In  Fig.  11  (H),  a  center  ash  "I"  is  stiffened  by  two 
spruce  side  plates,  the  ash  member  taking  the  bending 
moment,  and  the  spruce  the  compression.  Spar  (I)  has 
a  compound  central  "I,"  the  upper  and  lower  flanges  being 
of  ash  and  the  center  web  of  three  ply  veneer.  The  two 
outer  plates  are  of  spruce.  This  should  be  a  very  efficient 
section,  but  one  that  would  be  difficult  and  costly  to  build. 
Fig.  (J)  is  the  same  as  (B),  except  that  the  parting  lies 
in  a  horizontal  plane.  Spar  (K)  has  ash  top  and  bottom 
members,  and  spruce  or  veneer  side  plate.  The  resistance 


.      K  J  I  H 

Fig.    11.     BuiltUp    Wooden    Wing    Spars,    Commonly    Used    with    European 
Aeroplanes. 

of  this  shape  to  side  thrust  or  twist  would  be  very  slight. 
The  sides  are  both  screwed  and  glued  to  the  top  and 
bottom  members. 

The  front  end  of  a  Hansa-Brandenburg  wing  is  shown 
by  Fig.  12,  the  box  spar  and  its  installation  being  draw^n 
to  scale  and  with  dimensions  in  millimeters.  The  top  and 
bottom  are  sloped  in  agreement  with  the  rib  flange  curve, 
and  the  rib  web  is  strengthened  by  stiffeners  at  either 
side  of  the  spar.  The  hardwood  dowel  strips  are  at  top 
and  bottom  as  in  Fig.  B,  and  when  placed  in  this  posi- 
tion the  glued  joint  is  not  subjected  to  the  horizontal 
shear  forces.  The  walls  are  thicker  at  top  and  bottom  than 
at  the  sides,  in  order  to  resist  the  greater  vertical  forces. 


222 


WING  DETAILS 


For  the  same  reason  is  deeper  than  it  is  wide.  As  will 
be  remembered,  the  drag  is  very  much  less  than  the  lift, 
and  again,  the  drag  stress  is  greatly  reduced  by  the  in- 
ternal drag  wire  bracing. 


i-L. 


j-L ^. 


Ib 


^v 


Jli 


Leading  Edge  Construction.  In  the  early  Bleriot  mono- 
planes the  leading  edge  was  of  sheet  aluminum,  bent  into 
**U"  form  over  the  nose  of  the  rib.    In  modern  biplanes, 


WING  DETAILS 


223 


this  edge  is  generally  of  "U"  form  hollow  spruce,  about 
3/16  inch  thick.  Another  favorite  material  is  flattened 
steel  tubing,  about  l/2"xl/4",  and  of  very  light  gauge, 
the  long  side  being  horizontal.  The  tube  has  the  advan- 
tage of  being  much  thinner  and  much  stiffer  than  the 
other  forms,  and  the  thin  edge  makes  it  very  suitable  for 
certain  types  of  aerofoils.  The  wing  tip  bows  are  gen- 
erally of  hollowed  ash  and  are  fastened  to  the  spar  ends, 
leading  edges,  and  trailing  edges  with  maple  dowels,  the 
joint  being  of  a  long  scarfed  form.    When  the  scarfed 


Socnn 


iMmr 


Fig.  12b.  An  Old  Type  of  Curtiss  Biplane  Strut  Socket,  at  Left.  At  the 
Right  Is  a  More  Modern  Type  in  Which  the  Bolts  Do  Not  Pierce 
the  Spar. 


joint  is  doweled  together,  it  is  wrapped  with  one  or  two 
layers  of  glued  linen  tape.  In  some  types  of  machines 
the  top  surface  of  leading  edge  is  covered  with  thin  two 
ply  wood  from  the  extreme  front  edge  to  the  front  spar. 
This  maintains  the  aerofoil  curve  exactly  at  the  most 
critical  point  of  lifting,  and  also  stiffens  the  wing  against 
the  drag  forces. 

Trailing  Edges.  These  are  either  of  thin  beveled  ash  or 
steel  tube.  On  army  machines,  the  rear  part  of  the  trailing 
edge  fabric  is  pierced  with  holes  about  3/16  inch  diameter, 
the  holes  being  provided  with  rust  proof  eyelets.    This 


224 


WING  DETAILS 


relieves  an  excess  of  pressure  due  to  rips  or  tears ;  one 
opening  being  located  between  each  rib  and  next  to  the 
body. 

Protection  of  Wing  Wood  Work.  In  protecting  the 
wood  framework  of  the  wings  from  the  effects  of  moisture, 
at  least  three  coats  of  good  spar  varnish  should  be  care- 
fully applied,  with  an  extra  coat  over  the  glued  surfaces 
and  taping.  Shellac  is  not  suitable  for  this  purpose.  It 
cracks  with  the  deflection  of  the  wings  and  finally  admits 


Fig.  12c.  A  Standard  H-3  Interplane  Strut  Socket  Is  Shown  at  the  Left, 
the  Bolts  in  This  Case  Passing  on  Either  Side  of  the  Spar.  Note 
the  Stay  Wire  Attachment  Clips  and  Pinned  Strut  Connection. 
A  German  Strut  Socket  at  Right.     Courtesy  "Aerial  Age." 


water.  The  steel  parts  of  the  wing  should  be  given  two 
coats  of  fine  lead  paint,  and  then  two  coats  of  spar  varnish 
over  the  paint.  Wires  are  treated  with  some  flexible  com- 
pound, as  the  vibration  of  the  thin  wires,  or  cables, 
soon  cracks  ofT  any  ordinary  varnish.  The  use  of  shellac 
cannot  be  too  strongly  condemned ;  it  is  not  only  an  in- 
different protection,  but  it  causes  the  fabric  to  rot  when 
in  contact  with  the  doped  surface. 

Monoplane  Wing  Spars.    A  few  representative  mono- 
plane wing  spars  are  shown  by  Fig.   13,  the  R.  E.  P., 


WING  DETAILS 


225 


Bleriot  XI,  and  the  Nieuport.  Except  at  points  where 
the  stay  wires  are  connected,  the  Bleriot  spar  is  chan- 
neled out  into  "I"  beam  form  as  indicated  in  the  figure. 
It  will  be  noted  that  the  top  and  bottom  faces  of  the  spar 
are  slanted  to  agree  with  the  curvature  of  the  ribs.  A 
steel  connection  plate  is  bolted  to  the  sides  of  the  spar  by 
through  bolts,  and  with  a  lug  left  top  and  bottom  for  the 
top  and  bottom  guy  wires.  The  R.  E,  P.  is  also  an  'T" 
type,  the  section  "A"  being  taken  through  the  channeled 
portions,  while  *'B''  is  taken  through  one  of  the  connec- 


Fig.   13.     Typical  Monoplane  Wing  Spar  Construction. 

tion  points  where  the  beam  is  a  solid  rectangle.  The 
channeling  should  always  stop  at  connection  points ;  first, 
so  that  the  plates  have  a  good  bearing  surface,  and  second, 
to  allow  for  the  material  removed  by  the  bolt  holes. 

Probably  the  most  interesting  of  all  the  spars  is  the 
Nieuport,  which  is  a  combined  truss  and  girder  type.  This 
spar  tapers  down  from  the  center  to  both  ends,  being 
thickest  at  the  points  where  the  guys  are  attached.  The 
top  flange  (J),  and  the  bottom  flange  (L),  are  ash,  while 
the  side  plates  and  diagonals  (H)  are  spruce.  The  di- 
agonals resist  the  shear,  and  are  held  in  place  by  the  tie 


226  WING  DETAILS 

bolts  (I).  At  the  left,  the  spruce  cover  plate  is  removed, 
while  at  the  right  it  is  in  place  with  the  interior  construc- 
tion shown  in  dotted  lines.  The  dimensions  are  in  milli- 
meters. 

Location  of  Spars.   There  are  a  number  of  items  that 
affect  the  location  of  the  spars  in  regard  to  the  leading 


Fig.  14.  Plate  Connection  for  Monoplane  Stay  Wire  Connection  to  Spar. 
A  Compression  Member  or  Drag  Strut  Is  Sliown  in  the.  Center 
of  the  Spar  Which  Takes  Up  the  Thrust  Due  to  the  Angularity  of 
the  Stays  and  Also  the  Drag  Stress. 

edge.  The  most  important  factors  in  the  choice  of  loca- 
tion are:  (1)  Shape  and  depth  of  wing  section,  (2)  Center 
of  pressure  movement,  (3)  Drag  bracing  requirements, 
(4)  Width  of  ailerons,  (5)  Method  of  attaching  the  inter- 
plane  struts. 


CHAPTER  XI 
FUSELAGE  (BODY)  CONSTRUCTION. 

Purpose  of  Fuselage.  The  fuselage  of  a  monoplane 
or  tractor  biplane  is  the  backbone  of  the  machine.  It 
forms  a  means  of  connecting  the  tail  surfaces  to  the  main 
wing  surfaces,  carries  the  motor,  fuel  and  pilot,  and  trans- 
mits the  weight  of  these  items  to  the  wings  and  chassis. 
With  the  exception  of  the  wing  structure,  the  fuselage  is 
the  most  important  single  item  in  the  construction  of  the 
aeroplane.  Fig.  1  shows  a  typical  arrangement  of  a 
two-place  biplane  fuselage  equipped  with  a  water-cooled 
motor.  The  motor  E,  propeller  Y,  and  radiator  R  are 
placed  in  the  front  of  the  fuselage,  and  considerably  in 
advance  of  the  wings  W  and  D.  Immediately  behind 
the  motor  is  the  passenger's  seat  A  and  the  fuel  tank  F. 
The  pilot's  seat  B  is  placed  behind  the  trailing  edge 
of  the  wings  and  is  behind  the  passenger's  seat.  The 
cockpit  openings  G  and  H  are  cut  in  the  fuselage  top  for 
the  passenger  and  pilot  respectively.  The  rear  extension 
of  the  fuselage  carries  the  control  surfaces,  L  being  the 
vertical  fin,  M  the  rudder,  O  the  fixed  tail  or  stabilizer, 
and  P  the  elevator. 

Resistance.  To  reduce  the  resistance  in  flight,  the 
fuselage  is  of  as  perfect  streamline  form  as  possible,  the 
fuselage  being  deepest  at  a  point  about  one-third  from 
the  front.  From  this  point  it  tapers  out  gradually  to  the 
rear.  With  the  motors  now  in  use  it  is  only  possible  to 
approximate  the  ideal  streamhne  form  owing  to  the  front 
area  of  the  radiator  and  to  the  size  of  the  motor.  Again, 
the  projection  of  the  pilot's  head  above  the  fuselage  adds 

227 


228 


FUSELAGE  CONSTRUCTION 


considerably  to  the  resistance.  The  wind  shields  I  dis- 
turb the  flow  of  air.  The  connections  to  the  tail  surfaces 
and  to  the  chassis  members  K  also  add  to  the  total 
resistance.  An  arched  "turtle  deck"  J  is  generally  pro- 
vided, of  such  a  shape  that  the  pilot's  head  is  effectively 
"streamlined,"  the  taper  of  this  deck  allowing  the  dis- 


Figs.   1-2-3.     Fuselage  and  Motor  Arrangement  of  Tractor  Biplanes. 

turbed  air  to  close  in  gradually  at  the  rear.  The  flat 
area  presented  by  the  radiator  R  is  probably  the  great- 
est single  source  of  resistance,  and  for  this  reason  the 
radiator  is  sometimes  placed  at  the  side  of  the  fuselage, 
or  in  some  other  position  that  will  allow  of  a  better  front 
end  outline.  An  example  of  this  construction  is  shown 
by  Fig.  2  in  which  the  radiator  R  is  placed  behind  and 
above  the  motor  E.  The  front  fuselage  end  Z  can  now 
be  made  of  a  more  suitable  streamline  form. 


FUSELAGE  CONSTRUCTION  229 

Fig.  3  is  a  view  of  the  front  end  of  a  typical  fuselage 
in  which  an  air-cooled  rotary  type  of  motor  is  installed. 
Since  the  diameter  of  this  type  of  motor  is  seldom  much 
less  than  three  feet,  it  is  necessary  to  have  a  very  great 
diameter  in  the  extreme  front.  The  motor  housing  or 
"cowl"  marked  E  has  a  diameter  ''d"  which  should 
smoothly  blend  into  the  outline  of  the  fuselage  at  "b." 
In  the  older  types  of  construction  there  was  often  a  very 
considerable  break  in  the  outline  at  this  point,  especially 
in  cases  where  the  circular  cowl  was  abruptly  connected 
to  a  square  fuselage.  A  break  of  this  sort  greatly  in- 
creases the  head  resistance.  A  "spinner"  or  propeller 
cap  marked  Z  in  Fig.  3  is  an  aid  in  reducing  the  resistance 
offered  by  the  motor  cowl  and  also  reduces  the  resistance 
of  the  inner,  and  ineffective,  portion  of  the  propeller 
blades.  The  cap  in  any  case  is  smaller  than  the  cowl  open- 
ing in  order  that  cooling  air  be  admitted  to  the  enclosed 
cylinders. 

Distribution  of  Loads.  Returning  to  Fig.  1,  we  note 
that  the  weight  of  the  fuselage,  pilot,  passenger,  fuel,  con- 
trol surfaces  and  motor  are  carried  to  the  upper  wing  W 
by  the  "cabane"  strut  members  C-C  and  stays,  the  lower 
wing  being  connected  directly  into  the  sides  of  the  fusel- 
age. Continuations  of  the  cabane  members  on  the  inte- 
rior of  the  fuselage  inter-connect  the  upper  and  lower 
wings  (shown  in  dotted  lines).  The  interplane  stays  in 
connection  with  the  cabane  members  bind  the  wings  and 
fuselage  into  a  unit.  A  vertical  line  "CP"  passes  through 
the  center  of  pressure  of  both  wings,  and  approximately 
through  the  center  of  gravity  of  the  machine.  In  other 
words,  the  machine  is  nearly  balanced  on  the  center  of 
pressure  line.  The  turning  moments  of  weights  behind 
the  CP  must  approximately  balance  the  opposite  turning 
moments  of  the  masses  located  in  front  of  the  CP.  The 
exact  relation  between  the  center  of  pressure  and  the  cen- 
ter of  gravity  will  be  taken  up  later. 


230  FUSELAGE  CONSTRUCTION 

Variable  loads  such  as  the  passenger,  gasoline  and  oil, 
are  placed  as  nearly  as  possible  on  the  center  of  pressure 
line,  so  that  variation  in  the  weight  will  not  affect  the 
balance.  In  the  figure,  the  passenger's  seat  A,  and  the 
fuel  tank  F  are  on  the  CP  line,  or  nearly  so.  Thus,  a 
reduction  in  the  weight  of  the  fuel  will  not  affect  the 
''trim"  of  the  machine,  nor  will  a  wide  variation  in  the 
weight  of  the  passenger  produce  any  such  effect.  As 
shown,  the  passenger's  seat  is  placed  directly  on  the  top 
of  the  fuel  tank,  an  arrangement  widely  used  by  Euro- 
pean constructors.  In  the  majority  of  American  machines 
the  fuel  tank  is  placed  at  the  top  of  the  fuselage  instead 
of  in  the  position  illustrated.  As  the  pilot  is  considered 
as  a  constant  weight,  his  location  does  not  affect  the 
balance. 

When  at  rest  on  the  ground,  the  weight  of  the  rear 
end  of  the  fuselage  is  supported  by  the  tail  skid  N.  The 
length  of  this  skid  must  be  such  that  the  tail  surfaces 
are  kept  well  clear  of  the  ground.  The  center  of  the 
chassis  wheel  Q  is  placed  in  front  of  the  center  of  gravity 
so  that  the  weight  of  the  machine  will  cause  the  tail  skid 
to  bear  on  the  ground  when  the  machine  is  at  rest.  If  the 
wheel  were  behind  the  center  of  gravity,  the  machine 
would  "stand  on  its  nose"  when  making  a  landing.  The 
wheels  must  be  located  so  that  the  tendency  to  *'nose 
over"  is  as  small  as  possible,  and  yet  must  not  be  set  so  far 
forward  that  they  will  cause  an  excessive  load  on  the 
tail  skid.  With  too  much  load  on  the  skid,  the  tail  will 
not  come  up,  except  after  fast  and  prolonged  running, 
and  heavy  stresses  will  be  set  up  in  the  framework  due 
to  the  tail  bumping  over  the  ground  at  high  speed.  The 
skids  should  not  be  dragged  further  than  absolutely  nec- 
essary, especially  on  rough  ground.  With  proper  weight 
and  wheel  adjustment,  the  tail  should  come  up  in  a  short 
run.  The  wheel  adjustment  will  be  taken  up  under  the 
head  of  "Chassis." 


FUSELAGE  CONSTRUCTION  231 

Position  in  Flight.  In  normal  horizontal  flight,  the 
center  line  of  thrust  CT  is  horizontal  or  nearly  so.  This 
line  of  thrust  passes  through  the  center  of  the  motor 
crankshaft  and  propeller.  In  an  upward  climb,  the  CT 
is  inclined  at  the  angle  of  climb,  and  since  the  CT  indi- 
cates the  line  of  flight,  the  streamline  curves  of  the  body 
should  be  laid  out  so  that  the  axis  of  least  body  resistance 
will  coincide  with  the  line  of  thrust.  When  flying  horizon- 
tally at  the  normal  speed,  the  body  must  present  the  min- 
imum of  resistance  and  the  wings  must  be  at  the  most  effi- 
cient angle  of  incidence.  In  climbing,  or  flying  at  a  very 
low  speed,  the  tail  must  necessarily  be  depressed  to  gain 
a  large  angle  of  wing  incidence,  and  hence  the  body  re- 
sistance will  be  comparatively  high  owing  to  the  angle  of 
the  body  with  the  flight  line.  It  is  best  to  have  the  least 
resistance  of  the  fuselage  coincide  with  the  normal  hori- 
zontal flight  speed.  This  condition  at  once  establishes 
the  angle  of  the  wings  in  regard  to  the  fuselage  center 
line. 

Center  Line  of  Resistance.  The  center  line  of  thrust 
should  pass  through  the  center  of  total  head  resistance 
as  nearly  as  possible.  The  total  resistance  referred  to  is 
composed  of  the  wing  drag,  body  and  chassis  resistances. 
In  an  ordinary  miHtary  type  of  aeroplane  this  line  is  lo- 
cated approximately  at  one-third  of  the  gap  from  the 
bottom  wing.  Owing  to  variations  in  the  drag  of  the 
wings  at  different  angles,  this  point  varies  under  different 
flying  conditions,  and  again,  it  is  affected  by  the  form 
and  size  of  the  fuselage  and  chassis.  The  exact  location 
of  the  center  of  resistance  involves  the  computation  of  all 
of  the  resistance  producing  items. 

In  addition  to  passing  through  the  center  of  resistance, 
the  center  line  of  thrust  should  pass  slightly  below  the 
center  of  gravity  of  the  machine.  In  this  position  the  pull 
of  the  motor  tends  to  hold  the  head  up,  but  in  case  of 
motor  failure  the  machine  immediately  tends  to  head- 


-32  FUSELAGE  CONSTRUCTION 

dive  and  thus  to  increase  its  speed.  The  tendency  to  dive 
with  a  dead  motor  automatically  overcomes  the  tendency 
to  "stall"  or  to  lose  headway.  With  the  centerline  of 
thrust  determined,  and  with  given  motor  dimensions,  the 
fuselage  position  can  be  located  at  once  in  regard  to  the 
wings.  This  is  good  enough  for  a  preliminary  layout, 
but  must  be  modified  in  the  final  design.  As  before  ex- 
plained, the  centerline  of  thrust  is  located  at  a  point 
between  the  two  wings,  approximately  one-third  of  the 
gap  from  the  lower  wing.  In  machines  having  a  span  of 
35  feet  and  over,  it  is  a  trifle  less  than  one-third  the  gap, 
while  in  small  speed  scouts  it  is  generally  a  trifle  over 
one-third.  This  rule  checks  very  closely  with  the  data 
obtained  from  22  standard  machines.  Thus,  in  a  machine 
having  a  6-foot  gap,  the  propeller  centerline  will  be  located 
about  2  feet  above  the  lower  wing. 

The  top  of  the  fuselage  (measured  under  the  stabilizer 
surface)  is  from  5  to  8  inches  above  the  center  line  of 
thrust.  At  the  motor  end,  the  height  of  the  fuselage  above 
the  CT  is  controlled  by  a  number  of  factors,  either  by  the 
type  of  motor,  or  by  the  arrangements  made  for  access 
to  the  motor  parts.  In  a  number  of  European  machines, 
the  motor  sits  well  above  the  top  of  the  fuselage,  this 
always  being  the  case  when  a  six-cylinder,  vertical  water- 
cooled  motor  is  used.  With  an  air-cooled  type,  the  top 
is  governed  by  the  cowl  diameter. 

Motor  Compartment.  The  space  occupied  by  the  mo- 
tor and  its  accessories  is  known  as  the  "motor  compart- 
ment," and  in  monoplane  and  tractor  biplane  fuselage  it  is 
located  in  the  extreme  front  of  the  body.  The  interior 
arrangement  varies  with  different  types  of  motors  and 
makes  of  machines.  With  rotary  cylinder  motors,  the 
"compartment"  is  often  nothing  but  a  metal  cowl,  while 
with  large  water-cooled  motors  it  occupies  a  considerable 
portion  of  the  body.  Water-cooled  motors  are  generally 
covered  with  automobile  type  hoods,  these  usually  being 


FUSELAGE  CONSTRUCTIOX 


233 


provided  with  "gilled"  openings  for  ventilation.  Owing 
to  the  heat  generated  in  the  motor,  some  sort  of  ventila- 
tion is  imperative  at  this  point.  Whatever  the  type,  the 
compartment  is  always  cut  off  from  the  rest  of  the  fuselage 
by  a  fireproof  metal  partition  to  guard  against  fire  reach- 
ing the  passenger  or  fuel  tanks.  The  fuel  and  oil  should 
always  be  separated  from  the  motor  by  substantial  par- 
titions since  a  single  carbureter  "pop"  may  cause  serious 
trouble. 

Accessibility  is  a  most  important  feature  in  the  design 


Fig.  4.     Mounting    and     Cowls    for    Rotary     Cylinder     Motors. 
•Flight." 


Courtesy 


of  the  motor  end,  and  hence  the  hood  should  be  of  the 
hinged  automobile  type  so  that  it  can  be  easily  raised 
for  inspection  or  repairs.  In  the  Curtis  JN4-B  Military 
Training  Tractor,  the  cylinder  heads  and  valve  mechanism 
project  slightly  above  the  top  of  the  hood  so  that  these 
parts  are  amply  cooled  and  are  entirely  accessible.  Access 
to  the  carbureter  can  be  had  through  a  small  handhole  in 
the  side  of  the  hood.  The  radiator  in  this  Curtiss  model 
is  located  in  the  extreme  front  end  of  the  fuselage — auto- 
mobile fashion.  The  propeller  shaft  passes  through  a 
central  opening  in  the  radiator.  In  Fig.  6  the  vertical 
motor  E  is  set  down  low  in  the  frame,  the  upper  part  of 
the  fuselage  F  ending  at  H.    The  engine  bearer  B,  which 


23-i  FUSELAGE  CONSTRUCTION 

carries  the  motor,  forms  the  top  part  of  the  fuselage  at  this 
point.  The  engine  is  thus  in  the  clear  and  access  can 
easily  be  had  to  every  part  of  the  motor.  The  radiator 
is  in  front  of  the  motor  at  R.  When  in  flight  the  motor 
is  covered  by  a  sheet  metal  hood  similar  to  the  folding 
hood  used  on  automobiles.  This  type  is  used  in  the  Mar- 
tin, Sturtevant,  and  several  Eurpean  machines. 


Fig.  5.  Motor  Compartment  of  a  Curtiss  Tractor  Biplane  Using  a  Front  Type 
Radiator.  Note  the  Two  Exhaust  Pipes  Which  Carry  the  Gases 
Over  the  Top  of  the  Wings. 

Fig.  7  is  a  very  common  front  end  arrangement  used 
with  side  radiators.  The  top  fuselage  member  F  is 
brought  down  in  a  very  low  curve,  leaving  the  greater 
part  of  the  motor  projecting  above  the  fuselage.  At  the 
extreme  front,  the  upper  fuselage  member  F  joins  the 
engine  bearer  B,  the  connection  being  made  with  a  pressed 
steel  plate.  The  radiator  R  is  shown  at  the  side  of  the 
fuselage.     The  cylinders  are  not  usually  covered  when 


FUSELAGE  CONSTRUCTION 


235 


in  flight.    In  the  front  view  it  will  be  noted  that  the  radi- 
ators are  arranged  on  either  side  of  the  fuselage. 

A  side  view  of  the  H.  and  ]\1.  Farman  Fighter  is  shown 
by  Fig.  10.  This  a  very  efficient  French  machine  which 
has  seen  much  active  service  in  the  war.  The  front  end 
is  much  like  that  shown  in  Fig.  7  except  that  a  spinner 
cap  is  fitted  to  the  propeller  boss.  A  "V"  type  motor  al- 
lows of  the  radiator  being  mounted  between  the  two  rows 


Figs.  6-7,     Various    Motor   Arrangements,   and   Radiator   Locations. 

of  cylinders,  and  in  a  position  where  it  will  cause  the  least 
possible  head  resistance. 

The  fuselage  is  of  excellent  streamline  form  and  shows 
careful  study  in  regard  to  the  arrangement  of  the  power 
plant.  Unlike  the  majority  of  machines,  the  fuselage  is 
raised  well  above  the  bottom  wing,  this  being  done  evi- 
dently to  increase  the  range  of  the  gun  in  the  rear  cock- 
pit. The  increased  height  allows  the  gun  to  shoot  over 
the  top  plane  at  a  fairly  small  angle,  and  the  height  above 


FUSELAGE  CONSTRUCTION  237 

the  ground  permits  the  use  of  a  very  large  and  efficient 
propeller  without  having  an  excessively  high  chassis.  At 
the  rear  the  fuselage  tapers  down  to  a  very  thin  knife 
edge  and  therefore  produces  little  disturbance. 

Fig.  11  shows  a  Sturtevant  Training  Biplane  in  which 
the  radiator  is  mounted  at  the  front  edge  of  the  upper 
plane.  This  arrangement  was  originally  introduced  by 
the  Sturtevant  Company  in  their  steel  biplanes  and  has 
proved   a  very  efficient  type   for  cooling,   although   the 


Pig.  11.  Radiator  Mounted  at  Leading  Edge  of  Upper  Wing.  This  Type 
Is  Used  with  the  Sturtevant  and  Lawson  Aeroplanes  and  Is  Very 
Effective   Because   of  the   Improved   Circulation. 

radiator  must  affect  the  lift  of  the  top  plane  to  a  very 
considerable  extent. 

Pilot  and  Passenger  Compartments.  These  compart- 
ments contain  the  seats,  controls,  and  instruments,  and  in 
the  military  types  contain  the  gun  mounts  and  ammuni- 
tion. In  some  battleplanes,  the  passenger  or  "observer" 
occupies  the  rear  seat,  as  this  position  gives  a  wider  range 
of  fire  against  rear  or  side  attacks.  This  arrangement  is 
true  of  the  H.  and  M.  Farman  machine  just  illustrated 
and  described.  In  the  large  German  "Gotha"  the  gunner 
occupies  the  rear  position  and  fires  through,  or  above,  a 
tunnel  built  through  the  rear  end   of  the  fuselage.     In 


238  FUSELAGE  CONSTRUCTION 

some  forms  of  training  machines,  the  pilot  and  passenger 
sit  side  by  side  instead  of  in  tandem,  as  this  arrangement 
allows  better  communication  between  the  pilot  and  stu- 
dent, and  permits  the  former  to  keep  better  watch  over 
the  movements  of  the  student.  A  notable  example  of  this 
type  is  the  Burgess  Primary  Trainer.  A  side-by-side 
machine  must  have  a  very  wide  fuselage  and  therefore 
presents  more  head  resistance  than  one  with  the  seats 
arranged  in  tandem,  but  with  proper  attention  to  the 
streamline  form  this  can  be  reduced  so  that  the  loss  is 
not  as  serious  as  would  be  imagined  from  a  view  of  the 
layout. 

The  seats  may  be  of  several  types,  (a)  the  aluminum 
''bucket"  type  similar  to,  but  lighter  than,  the  bucket  seats 
used  in  racing  automobiles ;  (b)  the  woven  wicker  seat 
used  in  many  types  of  German  machines,  or  (c)  the  modi- 
fied chair  form  with  wooden  side  rails  and  tightly 
stretched  leather  back  and  bottom.  Whatever  the  type, 
they  should  be  made  as  comfortable  as  possible,  since 
the  operation  of  a  heavy  machine  is  trying  enough  without 
adding  additional  discomfort  in  the  form  of  flimsy  hard 
seats.  In  the  older  machines  the  seats  were  nothing  more 
than  perches  on  which  the  pilot  balanced  himself  precari- 
ously and  in  intense  discomfort.  A  few  pounds  added  in 
the  form  of  a  comfortable  seat  is  material  well  spent  since 
it  is  a  great  factor  in  the  efficient  operation  of  the  machine. 
Wicker  seats  are  light,  yielding  and  comfortable,  and  can 
be  made  as  strong  or  stronger  than  the  other  types.  It 
seems  strange  that  they  have  not  been  more  widely 
adopted  in  this  country. 

xA-ll  seats  should  be  slightly  tilted  back  so  that  the  pilot 
can  lean  back  in  a  comfortable  position,  with  a  certain 
portion  of  his  weight  against  the  back  of  the  seat.  Sit- 
ting in  a  rigid  vertical  position  is  very  tiring,  and  is  espe- 
cially so  when  flying  in  rough  weather,  or  on  long  recon- 
naissance trips.     The  backs  of  the  seats  should  be  high 


FUSELAGE  CONSTRUCTION 


239 


and  head  rests  should  be  provided  so  that  the  pilot's  head 
can  be  comfortably  supported  against  the  pressure  of  the 
wind.  If  these  head  rests  are  "streamUned"  by  a  long, 
tapering,  projecting  cone  running  back  along  the  top  of 
the  fuselage,  the  resistance  can  be  considerably  reduced. 
This  arrangement  was  first  introduced  in  the  Gordon- 
Bennett   Deperdussin   and   has   been   followed   in   many 


Fig.   12.     Hall-Scott  Motor  and  Side  Type  Radiator  Mounting  on  a  Typical 
Tractor. 

modern  machines.  In  the  Deperdussin,  the  pilot's  head 
was  exposed  directly  to  the  full  blast  of  the  propeller  slip- 
stream and  a  head  support  was  certainly  needed.  Small, 
transparent  wind  shields  are  now  fastened  to  the  front 
edge  of  the  cockpit  openings  which  to  a  certain  extent 
shield  the  pilot  from  the  terrific  wind  pressure.  These 
are  quite  low  and  present  little  resistance  at  high  speeds. 
A  heavy  leather  covered  pad,  or  roll,  should  be  run 
entirely  around  the  edge  of  the  cockpit  opening  as  a  pro- 
tection to  the  pilot  in  case  of  an  accident  or  hard  landing. 


240  FUSELAGE  CONSTRUCTION 

The  roll  should  be  at  least  3^"  in  diameter  and  should  be 
filled  with  horse  hair.  All  sharp  edges  in  the  cockpit 
should  be  similarly  guarded  so  that  in  the  event  of  the 
pilot  being  thrown  out  of  his  seat,  he  will  not  be  cut  or 
bruised.  Each  seat  should  be  provided  with  an  improved 
safety  strap  that  will  securely  hold  the  pilot  in  his  seat, 
and  yet  must  be  arranged  so  that  it  can  be  quickly  and 
easily  released  in  an  emergency.  In  flight  the  occupants 
must   be   securely   strapped   in   place   to   prevent  being 


Fig.  13.  Deperdussin  Monoplane  with  Monocoque  Body.  Note  the  Stream- 
line Form  of  the  Body  and  the  Spinner  Cap  at  the  Root  of  the 
Propeller. 

thrown  out  during  rapid  maneuvers  or  in  rough  weather. 
Buckles  should  be  substantial  and  well  sewed  and  riveted 
to  the  fabric  so  that  there  will  be  no  danger  of  their 
being  torn  out.  The  straps  must  be  arranged  so  that 
they  will  not  interfere  with  the  free  movement  of  the 
pilot,  and  so  that  they  will  not  become  entangled  with 
the  controls.  It  is  best  to  copy  the  sets  approved  by 
the  government  as  these  are  the  result  of  long  continued 
experiment  and  use. 

Care  must  be  taken  to  have  the  seats  located  at  the 
correct  height  from  the  floor  so  that  the  legs  will  not 
become  cramped.  In  the  majority  of  machines,  the  ver- 
tical rudder  is  operated  by  the  feet.     Unless  the  seat  is 


FUSELAGE  CONSTRUCTION  241 

at  the  proper  height,  the  pilot  will  be  in  a  strained  posi- 
tion as  he  cannot  shift  around  nor  take  his  feet  from  the 
rudder  bar.  Either  the  rudder  bar  or  the  seat  should 
be  adjustable  for  different  lengths  of  legs.  Usually  the 
adjustment  is  made  in  the  rudder  bar  since  it  is  not  usu- 
ally advisable  to  shift  the  seats  owing  to  the  necessity 
of  having  the  pilot's  weight  in  a  fixed  position.  In  some 
old  types  of  monoplanes,  the  pilot  sat  on  a  small  pad 
placed  on  the  floor  of  the  fuselage.  Needless  to  say,  this 
was  a  horribly  uncomfortable  position  to  be  in,  but  as  the 
flights  of  that  time  were  of  short  duration  it  did  not  mat- 
ter much.  If  the  feet  could  be  removed  occasionally  from 
the  rudder  bar  the  matter  of  seat  position  would  not  be 
of  so  much  importance,  but  to  sit  flat  on  the  floor,  with 
the  legs  straight  out,  for  a  couple  of  hours  is  a  terrible 
strain  and  has  undoubtedly  caused  many  accidents 
through  cramps. 

As  both  the  passenger  and  the  fuel  are  varying  weights, 
the  fuel  tank  seat  idea  is  good.  This  allows  both  of  these 
items  to  be  placed  at  the  center  of  gravity  of  the  machine 
where  weight  variation  will  have  no  effect  on  the  balance 
of  the  plane.  In  this  position,  however,  the  fuel  must  be 
pumped  up  to  a  higher  auxiliary  tank  since  the  main  tank 
would  be  too  far  below  the  carbureter  for  gravity  feed. 

The  flooring  of  the  cockpits  can  be  either  of  veneer, 
or  can  be  built  up  of  small  spruce  slats  about  3^"x^", 
the  slats  being  spaced  about  ^"  apart.  The  latter  floor  is 
specified  by  the  government  for  seaplane  use,  and  is  very 
light  and  desirable.  The  floor  is  placed  only  at  points 
where  it  will  be  stepped  on.  Observation  holes  are  cut 
in  the  floor  on  a  line  with  the  edge  of  the  seats  so  that 
the  occupants  can  view  the  ground  without  looking  over 
the  edge  of  the  fuselage.  The  observation  port  holes  are 
about  9  inches  in  diameter.  Glass  should  never  be  used 
in  the  cockpits  except  for  the  instrument  covers,  unless  it 
is  of  the  non-splintering  "triplex"  laminated  type  of  glass. 


2-12  FUSELAGE  CONSTRUCTION 

The  use  of  inflammable  celluloid  should  also  be  avoided 
as  being  even  more  dangerous  than  the  glass.  The  triplex 
glass  is  built  up  of  two  or  more  layers  of  glass,  which  are 
cemented  together  with  a  celluloid  film  applied  under 
heavy  pressure.  This  form  of  construction  is  very  strong, 
and  while  it  can  be  broken,  it  will  not  fly  apart  in  the  form 
of  splinters. 

All  instruments  should  be  placed  directly  in  front  of 
the  pilot  so  that  he  can  take  observations  without  turning 
his  head.  Usually  all  of  the  instuments,  with  the  excep- 
tion of  the  compass,  are  mounted  on  a  single  ''instrument 
board"  placed  in  front  of  the  pilot  and  directly  under  the 
forward  edge  of  the  cockpit  opening.  The  compass  can 
be  placed  on  the  floor  as  in  American  machines,  or  in- 
serted in  the  upper  wing  as  in  some  European  machines. 
The  motor  control  apparatus  is  placed  where  it  can  be 
reached  conveniently  from  the  seat.  Oil  gages,  gasoline 
gages,  and  revolution  counters  are  generally  placed  on 
the  instrument  board  where  they  can  be  easily  observed. 
If  a  wireless  set  is  carried,  the  switches  are  placed  on,  or 
near,  the  instrument  board.  Owing  to  the  uses  to  which 
the  different  machines  are  put  it  is  impossible  to  give  a 
list  of  instruments  that  would  be  suitable  for  every  ma- 
chine. The  simplest  machine  should  have  the  following 
instruments : 

Altimeter,  for  measuring  altitude. 

Clock  of  special  aeroplane  type. 

Incidence  indicator. 

Air  speed  indicators  for  measuring  the  speed  of  the 
machine  relative  to  the  air. 

Gasoline,  oil  and  pressure  gages  for  determining  amount 
of  fuel. 

Instruments  for  Navy  and  Army  machines  are  of  course 
more  complete.  In  the  specifications  for  Army  Hydro- 
aeroplanes (twin  motor  type  1916)  the  following  instru- 
ments are  specified: 


FUSELAGE  CONSTRUCTION 


243 


Aneroid  Barometer.  Graduated  in  feet,  and  reading  from 
sea-level  to  12,000  feet. 

Compasses.  One  in  each  cockpit.  To  be  of  the  Sperry 
Gyroscopic  type  with  an  elastic  suspension  and  prop- 
erly damped.  Shall  be  atached  to,  and  synchronized 
with,  the  ground  drift  indicator. 

Ground  Drift  Indicator.  Located  in  observer's  cockpit. 
For  noting  drift  due  to  side  winds.  See  illustration  on 
page  244. 

Clock.  Special  aeroplane  type,  built  to  resist  vibration. 
Located  in  pilot's  cockpit. 


Fig.    14.     Aeroplane  Compass  of  the   McCreagh-Osborn  Type.      (Sperry) 


Gasoline  Supply  Gage.  To  indicate  the  amount  of  fuel  in 
gravity  service  tank  at  all  attitudes  of  flight,  and  shall 
be  visible  from  pilot's  seat.  A  gage  in  the  main  tank 
will  also  be  desirable  that  will  register  the  approaching 
exhaustion  of  fuel.  This  indicator  should  register  when 
75%  of  the  fuel  in  the  main  tank  is  exhausted,  and  then 
record  the  remainder  continuously. 

Air  Speed  Indicator.     One  in  pilot's  cockpit. 

Angle  of  Incidence  Indicator.  Sperry  type.  To  be  located 
in  pilot's  cockpit. 

Inclinometer.    For  measuring  angle  of  inclination  of  Ion- 


244 


FUSELAGE  CONSTRUCTION 


gitudinal  axis  of  machine.    In  pilot's  cockpit,  and  placed 

on  instrument  board  in  the  vicinity  of  tachometers. 
Bank  Indicator.     For  indicating  the  proper  amount  of 

bank  on  turns.    In  pilot's  cockpit. 
Map  Board.    One  revolving  map  board  placed  in  pilot's 

cockpit. 
Map  Desk.     One  folding  map  desk  in  observer's  cockpit. 


Fig.   15.     Sperry   Ground  Drift  Indicator. 

Tachometers.    For  measuring  speed  of  motors  in  revolu- 
tions per  minute.    Pilot's  cockpit. 
Self-Starter  Switch.     For  operating  self-starter.     On  in- 
strument board  in  pilot's  cockpit. 

Among  the  other  accessories  specified  in  the  cockpit 
for  the  above  machines,  are  a  Pyrene  fire  extinguisher ; 
a  2-gallon  water  breaker ;  a  speaking  tube  for  communi- 
cation between  the  pilot  and  observer  (1"  to  1>^")>  ^  flash- 


FUSELAGE  CONSTRUCTION 


245 


light  signal  for  speaking  tube ;  and  a  tool  kit.    The  weight 
of  the  tool  kit  shall  not  exceed  11  lbs. 

General  Proportions  of  the  Fuselage.  The  total  length 
of  the  fuselage  depends  upon  the  type  of  power  plant, 
upon  the  span  or  chord  of  the  wings,  and  upon  the  ar- 
rangement of   the   tail  surfaces.     The  rear  end  of  the 


Fig.  16.  Cock-pit  of  a  "London  and  Provincial"  Biplane.  Control  Lever 
in  Foreground  and  Instrument  Board  Under  Cowl.  Courtesy  of 
•'Flight." 

fuselage  should  be  far  enough  away  from  the  wings  to 
insure  that  the  rear  surfaces  are  not  unduly  affected  by  the 
**down-wash"  or  the  "wake-stream"  of  the  wings.  A  very 
short  fuselage  gives  a  short  lever  arm  to  the  control 
surfaces,  hence  these  surfaces  must  be  very  large  with 
a  short  body.  With  the  stabilizer  surface  close  to  the 
wings,  the  "damping"  effect  is  slight,  that  is,  the  sur- 
face does  not  effectively  kill  or  "damp  down"  oscilla- 


246  FUSELAGE  CONSTRUCTION 

tions.  Large  tall  surfaces  are  heavy,  difficult  to  brace, 
and  cause  a  very  considerable  amount  of  head  resistance. 
The  extra  weight  of  a  long  fuselage  is  generally  offset  by 
the  increased  weight  caused  by  the  large  tail  area  of 
the  short  body  type. 

When  machines  are  crated  and  shipped  at  frequent  in- 
tervals, a  very  long  fuselage  is  objectionable  unless  it  is 
built  in  two  sections.  It  also  requires  much  storage 
space  and  a  very  long  hangar.  Machines  for  private  use 
must  often  be  sacrificed  from  the  efficiency  standpoint  in 
order  to  keep  the  dimensions  within  reasonable  limits. 
An  aeroplane  requiring  an  enormous  hangar  has  certainly 
no  attraction  for  the  average  man.  Every  effort  must  be 
made  to  condense  the  overall  dimensions  or  to  arrange 
the  extremities  so  that  they  can  be  easily  dismounted. 
Exhibition  flyers  require  specially  portable  machines  since 
they  ship  them  nearly  every  day,  and  the  expense  of  han- 
dling a  long  awkward  fuselage  may  alone  determine  the 
choice  of  a  plane.  It  is  usually  best  to  divide  the  body 
at  a  point  just  to  the  rear  of  the  pilot's  seat,  although 
many  flyers  look  upon  a  two-part  body  with  disfavor 
unless  they  can  be  shown  that  the  joint  connections  are 
as  strong  as  the  rest  of  the  fuselage. 

As  a  guide  in  the  choice  of  fuselage  proportions,  a  set 
of  diagrams  and  a  table  are  attached  which  gives  the  gen- 
eral overall  dimensions  of  several  prominent  makes  of 
machines.  The  letters  in  the  diagram  refer  to  the  letters 
heading  the  columns  in  the  tables  so  that  the  general 
dimensions  of  any  part  can  be  readily  determined.  I  do 
not  claim  that  these  dimensions  should  be  followed  relig- 
iously in  every  case,  but  they  show  what  has  been  done 
in  the  past  and  will  at  least  suggest  the  limits  within 
which  a  new  machine  can  be  built. 

Fig.  19  gives  the  outline  and  dimension  letters  for  two- 
place  machines  of  what  is  known  as  the  "Reconnaissance 
Type."     Both  water-cooled  and  air-cooled  motor  equip- 


FUSELAGE  CONSTRUCTION 


247 


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Fig.  19.  Fuselage  Dimension  Chart  for  Two  Place  Aeorplane  Fuselage. 
Upper  Diagram  Is  the  Water  Cooled  Type  and  the  Lower  Figure 
Applies  to  a  Machine  with  a  Rotating  Air  Cooled  Motor.  See  Table 
of  Dimensions  on  Page  248. 


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FUSELAGE  CONSTRUCTION  249 

ments  are  shown,  the  top  machine  being  of  the  water- 
cooled  type  while  the  lower  figure  shows  a  typical  two- 
place  machine  with  a  rotary  air-cooled  motor.  Under- 
neath this  side  elevation  is  a  front  view  of  the  fuselage, 
and  a  section  taken  through  the  point  of  greatest  depth. 
As  shown,  the  fuselage  is  of  square  cross-section,  but 
the  dimension  B  appHes  equally  to  the  diameter  of  a  cir- 
cular cross-section.  Dimension  C  gives  the  height  of  the 
curved  upper  deck,  or  "turtle  deck"  of  the  fuselage. 
Dimension  D  shows  the  extreme  length  extending  in 
front  of  the  leading  edge  of  the  lower  wing,  and  T  shows 
the  length  of  the  rear  portion  back  of  the  leading  edge  of 
the  lower  wing,  the  leading  edge  being  taken  as  a  base  of 
measurement.  The  location  of  the  deepest  section,  meas- 
ured from  the  extreme  front  of  the  fuselage,  is  given  by  E, 
the  depth  at  this  point  being  indicated  by  B.  The  extreme 
width  is  shown  by  I. 

In  the  machine  shown,  side  radiators  are  used,  the 
blunt  front  end  dimensioned  by  G  and  K  being  the  dimen- 
sion of  the  front  engine  plate.  When  front  radiators 
are  used,  the  dimensions  G  and  K  also  apply  to  the  size 
of  the  radiator.  The  amount  of  advance,  or  the  distance 
of  the  chassis  wheel  center  from  the  leading  edge  is  given 
by  S,  and  the  distance  of  the  wheel  center  below  the 
leading  edge  is  given  by  R.  V  is  the  length  of  the  engine 
projecting  above  the  fuselage  top.  The  passenger  or  ob- 
server is  indicated  by  1  and  the  pilot  by  2.  The  top 
plane  is  3  and  the  bottom  plane  4.  The  engine  is  located 
by  5,  and  the  top  fuselage-rail,  or  "longeron,"  by  7. 
Turtle  deck  is  6. 

FJg-  20  gives  the  diagrams  of  speed  scout  machines, 
both  of  the  air-cooled  and  water-cooled  types.  These  are 
the  small,  fast,  single  seat  machines  so  much  used  in  the 
European  war  for  repelling  air  attacks  and  for  guarding 
the  larger  and  slower  bombing  and  observation  machines. 
The  upper  drawing  shows  a  Curtiss  Speed  Scout  equipped 


250 


FUSELAGE  CONSTRUCTION 


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Fig.  20. 


Fig.  20.  Fuselage  Dimension  Diagrams  Giving  the  Principal  Dimensions  of 
Speed  Scout  Machines.  Upper  Figure  Shows  Typical  Scout  with 
Water  Cooled  Motor  (Curtiss),  While  Lower  Diagram  Shows  an 
Arrangement  Common   with  Rotary  Air  Cooled  Motors   (Nieuport) 


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252 


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with  a  "V"  type  water-cooled  motor  and  a  circulcT  front 
radiator.  While  the  front  of  the  body  is  circular,  it  grad- 
ually fades  out  into  a  square  cross-section  at  the  rear. 
The  lower  machine  is  a  Nieuport  speed  scout  equipped 
with  a  rotating  cylinder  air-cooled  motor.  In  this  scout, 
the  diameter  of  the  motor  cowl  is  given  by  dimension  K. 
Body  is  of  square  cross-section.  It  will  also  be  noted  that 
the  chord  of  the  lower  plane  is  less  than  that  of  the  upper 
plane  and  that  the  deep  body  almost  entirely  fills  the 
gap  between  the  two  wings. 


Fig.  21.     Curtiss  J  N  4-B  Fuselage  Boxed  for  Shipment. 


With  some  of  the  later  speed  scouts,  the  body  entirely 
fills  the  gap  between  the  wings  and  the  top  plane  is 
fastened  directly  to  the  top  members  of  the  fuselage. 
This  makes  windows  necessary  in  the  sides  of  the  fuselage. 
When  vertical  water-cooled  motors  are  used  on  speed 
scouts,  the  front  view  is  entirely  cut  off,  for  these  are 
very  large  motors  and  project  above  the  fuselage  for  a 
considerable  distance.  This  obstruction  is  avoided  in  the 
Curtiss  speed  scout  shown,  by  the  use  of  a  ''V"  type 
motor.  It  will  be  noted  that  these  two  scouts,  especially 
the  Nieuport,  are  of  excellent  stream  line  form,  a  very 
important  item  with  such  high  speed  machines.  The  pro- 
peller of  later  Nieuports  is  provided  with  a  conical  spinner 
•cap  which  evidently  reduces  the  head  resistance  to  a  con- 
siderable extent.  The  different  portions  of  the  machine 
are  indicated  by  the  same  figures  as  in  the  case  of  the 
reconnaissance  machine. 


CHAPTER  XII 
DETAILS  OF  FUSELAGE  CONSTRUCTION 

Classification  of  types.  While  there  are  a  number  of 
methods  adopted  in  building  up  the  fuselage  structure, 
the  common  type  is  the  *'wire  truss"  in  which  wood  com- 
pression members  are  used  in  connection  with  steel  wire 
or  cable  tension  members.  Four  wooden  "longerons" 
or  "longitudinals"  run  the  entire  length  of  the  body  and 
are  bent  to  its  general  outline.  The  longitudinals  are 
spaced  at  the  correct  distance  by  wood  compression  mem- 
bers, which  in  turn  are  held  in  place  by  wire  cross  brac- 
ing. This  method  of  trussing  forms  a  very  strong  and 
light  structure,  although  rather  complicated,  and  difficult 
to  build.  The  cross-section  is  rectangular,  although  in 
many  cases  the  body  is  converted  into  a  circular  or  ellip- 
tical section  by  the  use  of  light  wood  formers  fastened  to 
the  main  frame. 

Another  well  known  type  is  the  *'Monocoque"  body, 
first  used  on  the  Gordon-Bennett  Deperdussin  monoplane. 
This  fuselage  is  a  circular  shell  built  up  of  three-ply  tulip 
wood,  thus  forming  a  single  piece  body  of  great  strength. 
The  three-ply  shell  is  really  a  veneer,  the  layers  proceed- 
ing spirally  around  the  body,  each  layer  being  securely 
glued  to  its  neighbor.  Between  each  layer  is  a  scrim 
layer  of  treated  silk,  and  another  fabric  layer  is  generally 
glued  to  the  outside  of  the  shell.  The  shell  is  very  thin, 
the  total  thickness  of  the  three  layers  of  wood  and  fabric 
in  modern  machines  being  rather  less  than  1.5  millimeters 
(about  1/16  inch).  In  the  original  "Deps"  this  was  some- 
what greater,  0.15  inch.     Monocoque  construction  as  a 

253       . 


254  FUSELAGE  DETAILS 

rule  is  heavy  and  expensive,  but  offers  the  great  advantage 
of  strength,  perfect  alignment  at  all  times,  and  of  offering 
resistance  to  rifle  and  shell  fire.  If  the  longitudinals  of  a 
truss  type  are  struck  with  a  bullet,  or  shell  fragment,  the 
entire  fuselage  is  likely  to  fail,  but  a  monocoque  body 
may  be  well  perforated  before  failure  is  likely  to  take 
place. 

The  American  L.  W.  F.  Tractor  Biplane  has  a  mon- 
ocoque body  in  which  spruce  laminations  are  used  instead 
of  hardwood.  One  ply  runs  longitudinally  while  the  other 
two  layers  are  spiralled  to  the  right  and  left  respectively. 
Between  each  layer  is  a  scrim  layer  of  treated  silk,  the 
whole  construction  being  covered  with  a  final  layer  of 
fabric,  several  coats  of  waterproof  compound,  and  four 
final  coats  of  spar  varnish.  When  used  for  seaplanes  the 
wood  plies  are  stitched  together  with  strong  wires  to 
prevent  separation  due  to  dampness.  Since  spruce  is 
used  in  place  of  hardwood,  the  construction  is  lighter 
than  in  European  models,  and  the  L.  W.  F.  Company 
claim  that  it  is  lighter  than  the  usual  truss  construction. 
An  additional  advantage  of  the  monocoque  construction 
is  that  the  pilot  is  protected  against  splinters  or  penetra- 
tion by  the  limbs  of  trees  when  making  a  forced  landing  in 
the  brush. 

Another  form  of  monocoque  construction  was  adopted 
by  the  French  builder,  Bleriot,  at  the  beginning  of  the 
war.  The  fuselage  of  this  machine  was  covered  with 
papier-mache,  the  ash  longitudinals  being  buried  in  this, 
mixture.  The  papier-mache  is  built  up  with  glue  and  silk 
threads.  This  construction  is  very  light  and  strong,  but 
is  expensive  and  difficult  to  protect  against  moisture. 
The  front  of  the  fuselage  is  protected  with  a  3  millimeter 
steel  armor  plate  to  protect  the  pilot  against  bullets  and 
shrapnel.  The  papier-mache  portion  of  the  body  is  not 
easily  splintered  by  bullets. 

A  third  form  of  monocoque,  experimented  upon  by  the 


FUSELAGE  DETAILS  255 

author,  is  the  steel  shell  type  in  which  the  three-ply  wood 
veneer  is  supplanted  by  a  thin  steel  shell.  This  outer 
shell  is  strengthened  by  suitable  stiffener  angles.  With 
a  shell  thickness  of  0.013  inch,  the  strength  is  equal  to 
the  strength  of  a  wood  shell  and  is  slightly  less  in  weight. 
It  has  the  advantage  of  being  easily  and  cheaply  formed 
into  shape  and  is  absolutely  proof  against  the  influences 
of  heat  and  moisture.  It  cannot  splinter,  will  not  catch 
fire  and  offers  a  maximum  resistance  against  penetra- 
tion. There  is  yet  much  experimental  work  to  be  done 
before  the  construction  is  perfected. 

About  midway  between  the  truss  fuselage  and  the 
monocoque  is  the  veneer  construction  used  on  many  of 
the  modern  German  aeroplanes.  In  general,  this  may 
be  described  as  being  a  veneer  shell  fastened  to  the  con- 
ventional wood  longitudinals.  Stay  wires  are  not  in  gen- 
eral use,  the  veneer  taking  the  shear  due  to  the  bending 
movement.  Six  longerons  are  used  instead  of  four,  the 
two  additional  members  being  located  midway  on  the 
vertical  sides.  Transverse  wood  frames  take  the  place 
of  the  transverse  stay  wires  used  in  the  truss  type. 
Examples  of  this  type  are  met  with  in  the  "Albatros  de 
Chasse"  and  in  the  "Gotha"  bomb  dropper.  The  single 
seater,  '"Roland,"  has  a  fuselage  of  circular  section,  with  a 
true  monocoque  veneer  construction,  but  German-like, 
reinforces  the  construction  with  a  number  of  very  small 
longitudinals.  In  this  machine  there  are  6  layers,  or  plies, 
of  wood  reinforced  by  fabrics.  The  entire  thickness  of 
the  wood  and  fabric  is  only  1.5  millimeters  (1/16  inch). 

Steel  tube  fuselage  dates  back  to  the  beginning  of  the 
aeroplane  industry.  In  this  type  the  wood  longitudinals 
of  the  wood  truss  type  are  replaced  with  thin  gage  steel 
tubes,  the  cross  struts  being  also  of  this  material.  The 
diagonal  bracing  may  be  either  of  steel  wire,  as  in  the 
wood  frames,  or  may  be  made  up  of  inclined  steel  tube 
members  that  perform  both  the  duty  of  the  stay  wires  and 


256  FUSELAGE  DETAILS 

struts.  For  the  greatest  weight-efficiency,  a  steel  tube 
body  should  be  triangular  in  section  rather  than  square. 
A  triangular  section  saves  one  longitudinal  and  a  multi- 
tude of  wire  struts  and  connections  since  no  transverse 
bracing  is  necessary.  Connections  on  a  steel  tube  fuselage 
are  difficult  to  make  and  are  heavy.  They  require  much 
brazing  and  welding  with  the  result  that  the  strength  is 
uncertain  and  the  joint  is  heavy. 

A  very  modern  type  of  steel  construction  is  that  de- 
veloped by  the  Sturtevant  Company.  The  members  of 
the  Sturtevant  fuselage  are  in  the  form  of  steel  angles 
and  channels,  similar  in  many  respects  to  the  sections 
used  in  steel  buildings  and  bridges.  The  joints  are  riveted 
and  pinned  as  in  steel  structural  work.  The  longitudinals 
are  angles  and  the  struts  are  channels.  Crystallization 
of  the  steel  members  is  prevented  by  the  use  of  special  pin- 
connected  joints  provided  with  shock  absorbing  washers. 
Owing  to  the  simplicity  of  the  riveted  joints,  there  is 
practically  no  weight  due  to  connections,  and  since  the 
weight  of  connections  is  a  large  item  in  the  total  weight 
of  a  fuselage,  the  Sturtevant  is  a  very  light  structure. 
According  to  G.  C.  Loening,  engineer  of  the  company,  the 
fittings  of  a  large  wood  fuselage  weigh  at  least  60  pounds. 
This  is  almost  entirely  saved  with  the  riveted  connections. 

A  novel  type  of  wood  fuselage  has  been  described  by 
Poulsen  in  "Flight."  Eight  small  longitudinals  are  used 
Avhich  are  held  in  place  by  three-ply  wooden  formers  or 
diaphragms.  Wire  bracing  is  used  in  a  longitudinal  direc- 
tion, but  not  transversely  in  the  plane  of  the  diaphragms. 
The  cross-section  is  octagonal,  and  the  completed  struc- 
ture is  covered  with  fabric.  For  the  amateur  this  offers 
many  advantages  since  the  wiring  is  reduced  to  a  min- 
imum and  all  of  the  members  are  small  and  easily  bent  to 
shape.  It  is  fully  as  light  as  any  type  of  body,  for  the 
connections  are  only  thin  strips  of  steel  bolted  to  the 
diaphragms  with  small  machine  screws.    No  formers  are 


FUSELAGE  DETAILS 


257 


^^^^ 


258  FUSELAGE  DETAILS 

needed  for  the  deck,  and  the  machine  can  be  given  a  close 
approximation  to  the  ideal  stream-line  form  with  little 
trouble. 

Truss  Type  Fuselage.  We  will  now  take  up  the  con- 
struction of  the  truss  type  of  fuselage  in  more  detail,  and 
•investigate  the  merits  of  the  different  types  of  connec- 
tions used  in  fastening  the  frame  together.  Like  every 
part  of  the  aeroplane,  the  fuselage  must  either  be  right  or 
wrong,  there  is  no  middle  course.  Fig.  23  shows  a  side 
elevation  of  a  typical  truss  type  fuselage  built  up  with 
wood  longitudinals  and  struts,  the  tension  members  being 
high  tensile  strength  steel  wire  and  cable.  L  and  L'  are 
the  upper  and  lower  longitudinals,  S-S-S  are  the  vertical 
struts,  and  T-T-T  are  the  horizontal  cross  struts  which 
run  across  the  frame.  The  engine  bed  is  the  timber 
marked  B  at  the  front  of  the  body.  The  upper  wing  is 
attached  to  the  body  through  the  "cabane"  struts  C,  and 
the  chassis  connections  are  shown  at  D.  The  stern  post  E 
closes  the  rear  end  of  the  body  in  a  knife  edge  and  acts 
as  a  support  for  the  rudder  and  the  rear  end  of  the 
stabilizer.  F  is  the  seat  rail  which  carries  the  seats  and 
supports  the  control  yokes. 

All  cross  bracing  is  of  high  tensile  strength  steel  wire, 
or  of  high  strength  aviation  cable,  these  strands  taking  the 
tensile  stresses  while  the  wood  struts  are  in  compression. 
In  the  forward  portion,  double  stranded  cables  are  gen- 
erally used,  with  solid  wire  applied  to  the  after  portions. 
The  longitudinals  are  of  ash  from  the  motor  to  the  rear 
of  the  pilot's  seat,  while  the  rear  longitudinals  are  gen- 
erally of  spruce.  In  some  machines,  however,  the  entire 
length  of  the  longitudinals  is  ash.  The  latter  arrange- 
ment makes  a  heavier,  but  stronger  body.  The  struts  are 
usually  of  spruce  as  this  material  is  stififer  than  ash  and 
much  lighter. 

Both  the  struts  and  longitudinals  are  frequently  chan- 
nelled out  for  lightness,  as  shown  by  Fig.  27,  the  wooden 


260  FUSELAGE  DETAILS 

member  being  left  rectangular  in  section  only  at  the 
points  where  the  connections  are  made  with  the  struts 
and  cables.  The  channelling-out  process,  if  correctly  fol- 
lowed, gives  very  strong  stiff  members  with  a  minimum  of 
cross-sectional  area  and  weight.  Many  captured  German 
machines,  on  the  contrary,  have  solid  longitudinals  of 
rectangular  section,  wrapped  with  linen  fabric.  This  fabric 
strengthens  the  construction  and  at  the  same  time  re- 
duces the  chances  of  splintering  the  wooden  members 
in  a  hard  landing.  The  fabric  is  glued  to  the  wood  and 
the  entire  wrapping  is  then  given  several  coats  of  a  mois- 
ture repelling  varnish.  In  the  older  types  of  fuselage,  tlie 
longitudinals  were  often  of  the  "laminated"  class,  that  is, 
were  built  up  of  several  layers  of  wood  glued  together  in 
a  single  rectangular  mass.  This  reduced  the  tendency 
toward  splitting,  but  was  very  unreliable  because  of  the 
uncertainty  of  the  glued  joints  when  exposed  to  the  effects 
of  heat  and  moisture.  Laminated  longitudinals  are  now 
seldom  used,  particularly  in  the  region  of  the  motor  where 
water  and  oil  are  certain  to  wreck  havoc  with  the  glued  up 
members. 

As  the  stresses  rapidly  diminish  toward  the  tail,  it  is 
the  general  practice  to  taper  down  the  section  of  the 
longitudinal  toward  the  rear  and  to  reduce  the  section 
of  the  struts.  The  longitudinals  are  generally  kept  con- 
stant in  section  from  the  motor  to  the  rear  of  the  pilot's 
seat,  the  taper  starting  at  the  latter  point  and  continuing 
to  the  rear  end.  For  example,  if  the  longitudinal  section 
at  the  motor  is  l^"xl^",  the  section  at  the  rear  will  be 
V'xV\  the  width  of  the  struts  corresponding  to  this  taper. 
While  tapering  is  very  desirable  from  the  weight  stand- 
point, it  makes  the  fitting  problem  very  difficult  since  each 
fitting  must  be  of  a  different  dimension  unless  the  connec- 
tions can  be  designed  so  that  they  are  adjustable  to 
changes  in  the  section  of  the  longitudinals.  In  one  ma- 
chine, the  width  and  depth  of  the  longitudinals  are  kept 


FUSELAGE  DETAILS 


261 


constant,  the  variation  in  weight  and  section  being  accom- 
plished by  increasing  the  depth  of  the  channelling  as  the 
rear  is  approached.  With  this  design,  the  same  fittings 
can  be  applied  from  one  end  to  the  other. 

Since  the  loading  of  the  struts  is  comparatively  light, 
thev  can  be  much  reduced  in  section  by  channelling  or 


rrG.S>^^       FJ&2&B  ne.23-C^  -^^ 


Figs.  27-28-29-30.    Fuselage  Framing  Members  and  Details. 

by  chamfering,  as  shown  by  Fig.  28.  If  the  width  and 
thickness  is  maintained,  much  of  the  interior  material  can 
be  removed  without  danger  of  reducing  the  strength. 
Sketch  (A)  in  Fig.  28  shows  a  very  common  method  of 
strut  reduction,  the  strut  being  of  rectangular  section 
throughout  its  length,  but  tapered  in  such  a  way  that  it 
is  thickest  at  the  center  (d)  and  thinnest  at  the  two  ends 
(e).  To  obtain  the  correct  relation  between  the  center 
end  thickness  requires  very  careful  calculation.  As  shown. 


262  FUSELAGE  DETAILS 

the  strut  is  attached  to  the  upper  and  lower  longitudinals 
by  sheet  steel  fittings  or  "sockets."  Sketch  (B)  shows  a 
simple  method,  the  rectangular  strut  being  chamfered  off 
at  each  of  the  four  corners,  and  left  full  size  at  either  end 
where  the  fittings  connect  it  with  the  longitudinals.  This 
form  is  not  correct  from  a  technical  standpoint,  but  is 
generally  good  enough  for  lightly  loaded  struts,  and  has 
the  advantage  of  being  cheaply  and  easily  constructed. 
In  sketch  (C)  a  channelled  strut  is  shown,  the  center  por- 
tion being  channelled  out  in  a  manner  similar  to  the  chan- 
nelling of  the  longitudinals.  This  lightening  process  is 
most  commonly  adopted  with  the  large  heavily  loaded 
struts  in  the  front  portion  of  the  fuselage,  and  at  the 
points  where  the  motor  bed  is  suspended  or  where  the 
wings  and  chassis  are  attached  to  the  body.  The  black 
dots  at  the  ends  of  the  struts  indicate  the  bolt  holes  for 
the  fittings,  it  being  permissible  to  drill  holes  in  the  ends 
of  the  struts  but  not  in  the  longitudinal  members.  If  the 
strut  is  large  enough  to  resist  the  bending  stresses  at  the 
center  it  will  generally  allow  of  holes  being  drilled  near 
the  ends  without  danger  of  strength  reduction.  Again, 
the  struts  are  always  in  compression  and  hence  the  bolts 
may  be  depended  upon  to  partly  take  the  place  of  the 
removed  material  in  carrying  the  compressive  stresses. 
Holes  should  never  be  drilled  in  the  longitudinals  since 
these  members  may  be  either  in  tension  or  compression, 
depending  upon  the  angle  at  which  the  elevator  flaps  are 
set.  The  hole  not  only  destroys  the  strength  at  the  point 
at  which  it  is  drilled,  but  this  reduction  also  extends  to  a 
considerable  distance  on  either  side  of  the  hole,  owing  to 
the  fibrous  nature  of  the  wood.  In  steel  members  the 
effect  of  the  hole  is  purely  local  and  does  not  usually 
extend  much  beyond  the  edge  of  the  hole.  Considering 
the  wood  beam  as  consisting  of  a  series  of  parallel  fibers, 
it  will  be  seen  that  severing  any  one  of  the  fibers  will 
decrease  the  strength  of  the  wood  through  a  distance 


FUSELAGE  DETAILS  263 

equal  to  the  length  of  the  cut  fiber,  or  at  least  through  a 
distance  equal  to  the  natural  shear  value  of  the  resins  that 
bind  the  fibers  together. 

Fuselage  fittings  are  almost  numberless  in  the  variety 
of  design.  They  must  be  very  light  and  strong,  must  be 
applied  without  drilling  the  longerons,  and  should  be 
simple  and  cheap  to  construct.  They  are  usually  made 
of  sheet  steel  of  from  0.20  to  0.30  point  carbon,  and  may 
be  either  bent  or  pressed  into  shape.  At  the  points  where 
the  struts  are  joined  to  the  longitudinals,  the  fittings  con- 
nect struts  and  wires  in  three  planes,  the  vertical  struts 
and  fore  and  aft  wires  ;  the  transverse  wires  and  horizontal 
struts,  and  the  top  and  bottom  wires  that  lie  in  a  hori- 
zontal plane.  There  are  at  least  6  connections  at  every 
strut,  four  of  the  connections  being  made  to  the  stay 
wires  or  cables.  A  simple  connection  is  therefore  very 
hard  to  design. 

Fig.  29  shows  a  typical  fuselage  "panel"  and  the  inter- 
connected members  in  their  usual  relation.  LU  and  LL 
are  the  top  and  bottom  longitudinals  at  the  right,  while 
LU'  and  LL'  are  the  longitudinals  at  the  right  hand  side. 
The  vertical  struts  SV  and  SV  separate  the  top  and 
bottom  longitudinals,  while  the  horizontal  struts  SH  and 
SH'  separate  the  right  and  left  hand  sides  of  the  fuselage 
body.  The  wires  w-w-w-w  brace  the  body  fore  and  aft 
in  a  vertical  plane.  The  wires  t-t  lie  in  a  horizontal  plane, 
produce  compression  in  the  horizontal  struts  SH-SH',  and 
stiffen  the  frame  against  side  thrust.  The  transverse 
rectangle  SV-SH-SV'-SH'  is  held  in  shape  by  the  trans- 
verse stay  wires  W-W,  this  rectangle,  and  the  stays 
resisting  torsional  stress  (twisting),  act  against  the  struts 
composing  the  sides  of  the  rectangle.  In  some  European 
machines,  the  wires  WW  are  eliminated,  and  are  re- 
placed by  thin  veneer  panels,  or  short  wood  knee  braces 
as  shown  by  Fig.  30.  The  section  shows  the  longitudinals 
L-L-L-L  and  the  struts  SV-SV'-SH-SH'  braced  by  the 


264  FUSELAGE  DETAILS 

veneer  sheet  or  diaphragm  D.  This  diaphragm  is  well 
perforated  by  lightening  holes  and  effectually  resists  any 
torsional  stress  that  may  be  due  to  motor  torque,  etc. 
Since  the  transverse  wires  W-W  in  Fig.  29  are  rather 
inaccessible  and  difficult  to  adjust,  the  veneer  diaphragm 
in  Fig.  27  has  a  great  advantage.  In  this  regard  it  may 
be  stated  that  wire  bracing  is  not  a  desirable  construc- 
tion, and  the  substitution  of  solid  veneer  is  a  step  in 
advance. 

Wire  bracing  has  always  seemed  like  a  makeshift  to 
the  author.  The  compression  and  tension  members  being 
of  materials  of  widely  different  characteristics  are  not 
suitable  in  positions  where  a  strict  alignment  must  be 
maintained  under  different  conditions  of  temperature  and 
moisture.  The  difference  in  expansion  between  wire  and 
the  wood  compression  members  produces  alternate  tight- 
ness and  slackness  at  the  joints,  and  as  this  is  not  a  uni- 
form variation  at  the  different  joints,  the  frame  is  always 
weaving  in  and  out  of  line.  Under  the  influence  of  mois- 
ture the  wood  either  swells  or  contracts,  while  the  wire 
and  cable  maintain  their  original  lengths  and  adjustments. 
The  result  is  that  a  frame  of  this  kind  must  be  given 
constant  attention  if  correct  alignment  is  desired. 

The  adjustment  of  a  wire  braced  wood  fuselage  should 
be  performed  only  by  a  skilled  mechanic,  as  it  is  easily 
possible  to  strain  the  members  beyond  the  elastic  limit  by 
careless  or  ignorant  handling  of  the  wire  straining  turn- 
buckles.  In  the  endeavor  to  bring  an  old  warped  fuselage 
back  into  line  it  is  certain  that  the  initial  tension  in  the 
wires  can  be  made  greater  than  the  maximum  working 
stress  for  which  the  wires  were  originally  intended. 
Shrinkage  of  the  wood  also  loosens  the  bond  between  the 
wooden  members  and  the  steel  fittings  unless  this  is  con- 
tinually being  taken  up.  Some  form  of  unit  construction, 
such  as  the  monocoque  body,  is  far  more  desirable  than 
the  common  form  of  wire  trussed  wood  body. 


266  FUSELAGE  DETAILS 

Fuselage  Fittings.  In  the  early  days  of  aviation  the 
fuselage  fittings  on  many  machines  were  made  of  alu- 
minum alloy.  This  metal,  while  light,  was  uncertain  in 
regard  to  strength,  hence  the  use  of  the  alloy  was  grad- 
ually abandoned.  At  present  the  greater  part  of  the 
fittings  are  stamped  steel,  formed  out  of  the  sheet,  and 
are  of  a  uniform  strength  for  similar  designs  and  classes 
of  material. 

The  steel  best  adapted  for  the  fittings  has  a  carbon 
content  of  from  0.20  to  0.30,  with  an  ultimate  strength  of 
60,000  pounds  per  square  inch,  and  a  15  per  cent  elonga- 
tion. The  steel  as  received  from  the  mill  should  be 
annealed  before  stamping  or  forming  to  avoid  fracture. 
After  the  forming  it  can  be  given  a  strengthening  heat 
treatment.  A  lower  steel  lying  between  0.10  and  0.15 
carbon  is  softer  and  can  be  formed  without  annealing 
before  the  forming  process.  This  material  is  very  weak, 
however,  the  tensile  strength  being  about  40,000  pounds 
per  square  inch.  Fittings  made  of  the  0.15  carbon  steel 
will  therefore  be  heavier  than  with  the  0.30  carbon  steel 
for  the  same  strength.  The  thickness  of  the  metal  will 
vary  from  1/32^'  to  1/16",  depending  upon  the  load  com- 
ing on  the  fitting. 

A  typical  fuselage  strut  fitting  is  shown  by  Fig.  31-A 
in  which  L-L-L  are  the  longerons,  d  is  the  fitting  strap 
passing  over  the  longerons,  S  and  S'  are  the  vertical 
and  horizontal  struts  respectively.  The  stay  wires  are 
fastened  to  ears  (b)  bent  out  of  the  fitting,  the  wires 
being  attached  through  the  adjustable  turnbuckles  (t). 
The  struts  are  provided  with  the  sheet  steel  ferrules 
marked  (F).  There  are  no  bolts  passing  through  the 
longitudinals  L-L',  the  fitting  being  clamped  to  the 
wooden  member.  This  is  very  simple  and  light  fitting. 
Fig.  31-B  is  a  similar  type,  so  simple  that  further  dis- 
cussion is  unnecessary. 

Fig.  32  shows  a  fuselage  strut  fitting  as  used  on  the 


FUSELAGE  DETAILS 


267 


Standard  Type  H-3  Biplane.  We  are  indebted  to  "Aerial 
Age"  for  this  illustration.  This  consists  of  a  sheet  metal 
strap  of  "U"  form  which  is  bent  over  the  longitudinal  and 
is  bolted  to  the  vertical  strut.  At  either  side  of  the  strut 
are  through  bolts  to  which  bent  straps  attach  the  turn- 
buckles.  These  straps  are  looped  around  the  bolts  and 
form  a  clevis  for  the  male  ends  of  the  turnbuckles. 

An  old  form  of  fuselage  connection  used  on  the  Nieuport 


n^.airs 


Fig.   31.     Typical  Fuselage  Strut  Fittings. 

monoplane  is  shown  by  Fig.  33,  an  example  of  a  type  in 
which  the  bolts  are  passed  through  the  longeron  member. 
This  fitting  is  very  light  but  objectionable  because  of  the 
piercing  of  the  longeron. 

An  Austrian  aeroplane,  the  Hansa-Brandenberg,  has  a 
wood  fuselage  in  which  no  stay  wires  are  used.  This 
fuselage  is  shown  by  Fig.  23a.  Both  the  vertical  and 
inclined  members  are  wood  struts.  The  outer  covering 
of  wood  veneer  makes  the  use  of  stay  wires  unnecessary 
since  the  sheath  takes  up  all  horizontal  stresses,  and  hence 


268 


FUSELAGE  DETAILS 


forms  a  sort  of  plate  girder  construction.  The  German 
Albatros  also  employs  a  wireless  veneer  fuselage,  the  con- 
struction being  shown  in  detail  by  Figs.  36  and  36a.  Three 
longerons  are  located  on  either  side  of  the  body,  the  third 
member  being  placed  at  about  the  center  of  the  vertical 
side.    As  will  be  seen,  the  veneer  makes  the  use  of  wire 


Fig.   32.     Fuselage    Strut    Fittings    of   the    Standard   H-3    Training    Biplane. 

bracing  and  metal  connections  unnecessary.    The  veneer 
also  insures  perfect  alignment. 

Wing  Connections.  The  lower  wings  are  attached  to 
the  low^er  longitudinals  by  a  special  sheet  steel  fitting 
which  also  generally  connects  to  a  vertical  strut  at  this 
point,  and  to  an  extra  heavy  horizontal  strut.  A  sheet 
metal  clevis,  or  socket,  on  the  wing  spar  is  pinned  to  the 
fuselage  half  of  the  fitting  so  that  the  wing  can  be  easily 
detached  when  the  machine  is  to  be  dissembled.  At  this 
point  a  connection  is  also  provided  for  the  end  of  the  inner 
interplane  stay  wires.    The  horizontal  strut  at  the  point 


n 


US! 


cCZ 


'  r 


L 


FUSELAGE  DETAILS 


269 


of  wing  attachment  is  really  a  continuation  of  the  wing 
spar  and  takes  up  the  thrust  due  to  the  inclination  of  the 
interplane  stays.  In  the  majority  of  cases  the  horizontal 
thrust  strut  is  a  steel  tube,  with  the  hinged  connection 
brazed  to  its  outer  ends.  This  is  one  of  the  most  im- 
portant and  heavily  loaded  connections  on  the  machine 
and  should  be  designed  accordingly. 

Fig.  Z7  shows  a  typical  wing  to  fuselage  connection  of 
the  hinge  type.  The  wing  spar  (G)  is  covered  with  a 
sheet  steel  ferrule  (A)  at  its  inner  end.  Two  eye  bars 
(B)  are  bolted  to  the  wing  spar,  and  over  the  ferrule,  the 
eyes  of  the  bar  projecting  beyond  the  end  of  the  spar. 


Fig.  II.     Fuselage    Fittings    of    the    Nieuport    Monoplane. 


This  forms  the  wing  half  of  the  connecting  hinge.  The 
eyes  are  fastened  to  the  fuselage  hinge  member  (H)  by 
means  of  the  pin  (E).  This  pin  has  a  tapered  end  for  easy 
entry  into  the  joint,  and  is  pierced  with  holes  at  the  outer 
end  for  cotter  pins  or  a  similar  retaining  device.  The 
fuselage  hinge  member  (H)  is  brazed  to  the  end  of  the 
steel  tube  strut  (T).  This  tube  runs  across  the  fuselage 
from  wing  spar  end  to  wing  spar  end. 

Strut  tube  (T)  lies  on,  and  is  fastened  to,  the  fuselage 
longeron  (L),  and  also  lies  between  the  two  halves  of  the 
vertical  strut  (S).  The  vertical  strut  is  cut  out  at  its 
lower  end  for  the  receipt  of  the  steel  tube  (T).  A  steel 
plate  is  brazed  to  the  tube,  is  wrapped  about  the  longeron 
(L)  and  is  bolted  to  the  vertical  strut  (S).  The  interplane 
stay  (F)  is  attached  to  the  pin  (E)  at  the  point  of  junc- 
ture of  the  wing  spar  eye  and  the  fuselage  member  of  the 


270 


FUSELAGE  DETAILS 


-it 

1' 

£nvergure    sup?     6*60 

'•      "^       0  • 

R.f<(xlu,id    f,om     -VAtrofhl,-     from 

i 

drawtufs  madi  by  M    LtgOTgttti,  vnlh 

the  assislanci  ef  Ikt  Siniu  d  Aviation 

Jk 

M,Ulairt 

The  "D.I."  itpt  "Albatroj  de  Ch»M«," 

or  destroyer. 

Wings   witbont    V   or   .rrow.  aod   •: 

,'^r^ 

almost  eqaal  fpio. 

.j.ii 

Single  piece   upper  wing 

\ 

Single   piece  elevator. 

•\ 

Thick-coded    luielige 

\ 

160170  b.p.  Mercedes  engine 

^ 

Twin  macbine-gons   (marked   "Mitr.") 

L..., 

on  each  side  of  cngide  in  plan  view. 

The  tecllan  ot  fuselage  Ihroogh  Jnncllon 
•I  raddcr  po»t  it  ibown  en  lelt  ot  lalL 


Fig.  36.  Veneer  Fuselage  Construction  of  the  German  "Albatros"  Speed 
Scout.  Body  Outline  Is  Obtained  by  Veneer  Diaphragms  and  no 
Stay  Wires  Are  Used. 


FUSELAGE  DETAILS 


271 


hinge.  A  collar  (I)  Is  brazed  to  the  tube,  and  forms  a 
means  of  attaching  the  fuselage  stays  (D).  The  drift 
wires  (C)  of  the  wings  are  attached  to  an  eye  at  the  end 
of  one  of  the  wing  spar  bolts.  As  shown,  the  fitting  (H) 
is  a  steel  forging,  very  carefully  machined  and  reduced  in 
weight.  The  inside  wing  ribs  are  indicated  by  (K),  from 
which  it  will  be  seen  that  there  is  a  gap  between  the  end 
of  the  wings  and  the  outside  face  of  the  fuselage. 

Fig.  36-a  shows  the  construction  of  the  wing  joint  of 
the  German  Albatros  machine.  The  fuselage  is  of  mon- 
ocoque  construction  which  allows  of  a  simple  attachment 


Th«  AIb«tpo«  D.  I— Th« 
attachment  of  the  eabant  to 
tbm  fuMlage,  •howing  tb« 
adjustment  for  altera- 
ttpa  of  atagfter. 


The  sockets  securing  tfae  chassis 
struts  to  the  futelage.  The  stmts 
may  quickly  be  detached  by 
loosening  the  bolts  on  the  sockets. 


The  futelage  constmctioa. 
Fig.  36-a.    Details  of  Albatros  Veneer  Fuselage  Construction. 

to  the  outer  shell.  This  is  a  very  sturdy  and  simple  con- 
nection. Fig.  38-Z  is  the  wing  attachment  detail  of  the 
English  London  and  Provincial  Biplane  (1916),  the 
fuselage  in  this  case  being  of  the  wire  trussed  wood  type. 
We  are  indebted  to  ''Flight"  for  this  illustration. 

In  some  machines  the  interplane  stay  wires  are  attached 
to  a  lug  formed  from  the  attachment  plate,  but  we  do  not 
consider  that  this  construction  is  as  good  as  the  type  in 
which  the  wire  is  attached  directly  to  the  wing  spar  pin. 
While  the  former  may  be  easier  to  assemble,  the  attach- 
ment of  the  wire  to  the  pin  eliminates  any  eccentricity,  or 
bending  moment,  due  to  the  pull  of  the  interplane  stay. 
The  attachment  in  the  L.  W.  F.  insures  against  any  eccen- 


272 


FUSELAGE  DETAILS 


tricity  in  the  stay  attachment,  and  at  the  same  time  makes 
the  assembly  and  dismounting  a  very  simple  matter. 

Chassis  Member  Attachment.  The  attachment  of  the 
chassis  struts  generally  involves  some  difficulty  as  these 
members  usually  intersect  the  line  of  the  longerons  at  a 
very  awkward  angle.  If  the  wing  attachment  is  near  the 
same  point,  as  it  generally  is,  the  detail  is  made  doubly 


Fig.   n.     Wing    Connection    to    Fuselage. 

difficult.  The  chassis  must  be  pin  connected  as  in  the 
case  of  the  wing  joint  so  that  the  chassis  members  can  be 
easily  and  quickly  removed.  A  detail  of  a  chassis  to  body 
connection  is  shown  by  Fig.  39.  In  this  figure  (L)  is  the 
lower  longeron,  (S)  is  the  vertical  fuselage  strut,  and  (C) 
is  one  of  the  chassis  members.  The  upper  end  of  the 
chassis  member  is  enveloped  in  a  sheet  steel  ferrule  (D) 
which  is  bolted  in  place,  and  which  is  provided  with  a 
clevis  at  its  upper  end  for  the  attachment  pin  (P). 

A  plate  (E)  is  bolted  to  the  fuselage  strut  (S)  and  is 
passed  around  the  lower  longeron  (L),  a  hinge  joint  (H) 


FUSELAGE  DETAILS 


273 


being  provided  for  attachment  to  the  chassis  ferrule 
through  the  pin  (P).  Ears  or  lugs  are  left  at  (G-G)  for 
the  attachment  of  the  fuselage  stays  (B-B).  On  the  inner 
side  of  the  plate  (E)  are  attachment  lugs  for  the  hori- 
zontal strut  (H).  It  will  be  noted  that  the  plate  (E) 
is  well  provided  with  lightening  holes  so  that  the  weight 
can  be  kept  down  to  a  minimum.    The  pin  is  tapered  at 


rigY 


Fig.  38x.  Wing  Connection  of  the  Albatros  Reconnaissance  Biplane.  Fi( 
38y.  Wing  Attachment  of  Albatros  Fighter  with  Pin  Joint.  Fi{ 
38z.      Wing  Connection   of  London   and   Provincial   Biplane. 


the  end,  and  is  provided  with  cotter  pin  holes.  The  fitting 
in  general  is  small,  and  does  not  produce  any  great  degree 
of  head  resistance,  the  small  part  exposed  being  of  good 
streamline  form. 

Great  care  should  be  taken  in  brazing  or  welding  these 
fittings,  since  the  heat  changes  the  structure  of  the  metal 
and  greatly  reduces  its  strength.     The  brazing  tempera- 


274 


FUSELAGE  DETAILS 


ture  varies  from  1,500  to  1,700  degrees,  a  point  well  above 
the  tempering  heat  of  steel.  Attempts  have  been  made  to 
heat  treat  the  metal  after  the  brazing  operation,  but  with 
very  little  success,  owing  to  the  fact  that  the  heat  treat- 
ing temperature  is  generally  at  or  above  the  melting  point 
of  the  brazing  spelter,  hence  is  likely  to  cause  holes  and 
openings  in  the  brazed  joints.  With  acetylene  welded 
joints  the  parts  can,  and  should  be,  heat  treated  after  the 
welding.    While  this  is  an  apparent  advantage  of  acety- 


Fig.  39.    Chassis  Connection. 

lene  welding,  all  parts  cannot  be  successfully  handled  in 
this  manner.  The  welding  torch  can  only  join  edges, 
while  the  brazing  spelter  can  be  appHed  over  almost  any 
area  of  surface.  Welding  is  very  successful  in  joining 
thin  steel  tubes  while  in  many  fittings  made  of  sheet  metal, 
brazing  is  the  only  feasible  operation. 

Both  methods  have  a  common  fault,  in  that  they  are 
unreliable.  Imperfect  welds  and  brazing  are  not  always 
apparent  from  the  outside,  actual  breakage  of  the  part 
being  necessary  to  determine  the  true  nature  of  the  joint. 


FUSELAGE  DETAILS 


275 


FUSELAGE    WEIGHTS. 

Distribution  of  Weight.  The  weight  of  a  fuselage  de- 
pends upon  the  span  of  the  wings,  upon  the  seating 
capacity,  and  upon  the  weight  and  type  of  the  power 
plant.  The  weight  also  varies  considerably  with  the 
type  of  construction,  that  is,  whether  of  truss,  veneer, 
or  monocoque  construction.  A  heavily  powered  machine, 
or  one  carrying  more  than  a  single  person,  requires 
heavier  structural  members  and  hence  weighs  more  than 
a  small  single  seater.  The  amount  of  fuel  carried  also  has 
a  considerable  bearing  on  the  fuselage  weight. 

Probably  the  best  method  of  treating  this  subject  is  to 
give  the  fuselage  weights  of  several  types  of  well  known 
machines.  The  reader  will  then  have  at  least  a  com- 
parative basis  for  determining  the  approximate  weight. 
(Truss  type  only.) 

FUSELAGE    WEIGHTS 


Name 

Year 
Type 

Length 
Total 

H.  P. 

Motor 

Span   of 
Wings 

No.  of 
Seats 

Weight 
Bare 

Wt.   Fully 

Equipped ; 

.'anks,  Pipe 

Etc. 

Wt..  Fully 

Equipped, 

Including 

Power  Plant 

Aviatik 
( Tighter » 

1916 

21'  6" 

170 
Benz 

T=  40'  8" 
L=35'  5" 

2 

757.0 

1527 

Standard 
Reconn. 

1917 
H-3 

24'  6" 

135 
H-S 

T  =  40'  1" 
L  =  40'  1" 

2 

302.0 

479.7 

1218.5 

Curtiss 
School 

1917 
JN-4 

24'  6" 

90 
OX 

T  =  43'  7" 
L  =  33'  11" 

2 

296.3 

408.9 

904.7 

Nieuport 
Scout 

1916 
"11" 

15'  4" 

80 
LeR. 

T  =  24'  8" 
L  =  24 '  3" 

1 

372.0 

583.0 

There  are  so  many  variables  that  the  weight  cannot  be 
determined  by  any  set  rule  or  formula.  Alexander 
Klemin  in  his  "Course  in  Aerodynamics  and  Airplane 
Design"  says  that  the  approximate  weight  of  a  bare 
wood  truss  type  fuselage  is  about  150  pounds  for  a  ma- 
chine having  a  total  weight  of  2,500  pounds.  For  small 
biplane  and  monoplane  scouts  weighing  approximately 
1,200  pounds  total,  the  bare  fuselage  frame  will  weigh 
about  70  pounds.  These  figures  are  for  the  bare  frame 
alone  and  without  seats,  controls,  tail  skids  or  other  fit- 


276  FUSELAGE  DETAILS 

tings.  The  weights  given  under  the  column  headed  "Wt, 
Bare"  include  the  engine  beds,  tail  skids,  flooring,  cowling 
and  body  covering,  and  hence  exceed  the  *'bone  bare"  esti- 
mate of  Klemin  by  a  considerable  amount. 

The  all-steel  fuselage  of  the  large  Sturtevant  battle- 
plane (Model  A)  weighs  165  pounds  inclusive  of  the  steel 
engine  bed.  A  wooden,  wire  braced  fuselage  of  the  same 
size  and  strength  weighs  well  over  200  pounds,  the  metal 
fittings  and  wires  weighing  about  60  pounds  alone.  Ash 
is  used  in  the  wood  example  for  the  longerons.  The  struts 
and  diagonal  members  in  the  Sturtevant  metal  fuselage  are 
riveted  directly  to  the  longitudinals,  without  fittings  or 
connection  plates.  The  safety  factor  for  air  loads  is  8,  and 
for  the  ground  loads  due  to  taxi-ing  over  the  ground,  a 
safety  factor  of  4  is  used. 

After  a  minute  comparison  of  the  items  comprising  the 
fuselage  of  the  Curtiss  JX4-B  and  the  Standard  H-3, 
Klemin  finds  that  the  fuselage  assembly  of  the  Standard 
H-3  amounts  to  13.6  per  cent  of  the  total  loaded  weight, 
and  that  the  fuselage  of  the  Curtiss  JN4-B  is  15.5  per  cent 
of  the  total.  Tanks,  piping  and  controls  are  omitted  in 
both  cases.  For  machines  weighing  about  2,500  pounds, 
Dr.  J.  C.  Hunsaker  finds  the  body  weight  averaging  8.2 
per  cent  of  the  total,  this  figure  being  the  average  taken 
from  a  number  of  machines. 

On  careful  examination  it  will  be  found  that  the 
fuselage  assembly  (bare)  amounts  to  a  trifle  less  than  the 
wing  group  for  biplanes  having  a  total  weight  of  from 
1,900  to  2,500  pounds.  The  relation  between  the  wing 
weight  and  the  fuselage  weight  seems  to  bear  a  closer  rela- 
tion than  between  the  fuselage  and  total  weights.  We 
will  set  these  different  relations  forth  in  the  following 
table : 


FUSELAGE  DETAILS 


277 


PERCENTAGE    OF    FUSELAGE    WEIGHT 


Name    of    Plane    or 
Investigator 

Fuselage   Weight   as    Percentage   of   the 
Total  Load 

Wing 
Weight 
As    Per- 
centage of 

Total 
Weight 

Body 

Assembly 

Bare 

Body 
Assembly 

and 
Equipment 

Body 
Assembly 
and  Power 

Plant 

Curtiss  JX4-B 

I5.50<7f 

17.869-0 

43.96rf 

14.15% 

Standard   H-3 

13.60  Tr 

17.707O 

45.40  <r<r 

14.52% 

J.   C.   Hunsaker 

8.20 '-f 

11.50% 

34.309r 

16.50% 

Author's     Experience 

14.96^0 

17.6295: 

46.66% 

14.60  9'<- 

Average    of   Above 

13.06^r 

16.179'r 

42.589^ 

14.94% 

In  the  above  table,  the  column  headed  "Body  Assembly 
and  Equipment"  includes  the  body  frame,  controls,  tanks 
and  piping.  In  the  fourth  column,  the  radiator,  motor, 
propeller,  water,  and  exhaust  pipe  have  been  added.  For 
the  average  value  it  will  be  seen  that  the  bare  fuselage  is 
about  1.88  per  cent  lower  than  the  weight  of  the  wings. 
It  should  be  noted  that  the  wing  weight  given  is  the 
weight  of  the  surfaces  alone,  and  does  not  include  the 
weight  of  the  interplane  struts,  w^ires  and  fittings.  The 
weight  of  the  wing  surfaces  as  above  will  average  about 
0.75  pounds  per  square  foot. 

Based  on  the  above  figures,  we  can  obtain  a  rough  rule 
for  obtaining  the  approximate  weight  of  the  fuselage,  at 
least  accurate  enough  for  a  preliminary  estimate.  If 
A  =  the  total  area  of  the  wings,  then  the  total  weight 
of  the  wings  will  be  expressed  by  w  =  0.75A.  The  weight 
(f)  of  the  fuselage  can  be  shown  as  f=  13.06/14.94  x  w 
=  0.65  A. 

Example.  The  area  of  the  Standard  H-4  is  542  square 
feet  total.  Find  the  approximate  weight  of  the  fuselage. 
By  the  formula,  f  =  0.65 A  =  0.65x542  =  352  pounds.  The 
actual  bare  weight  is  302.0  pounds.  For  several  other 
machines,  the  actual  weight  is  greater  than  the  weight 
calculated  by  the  formula,  so  that  the  rule  can  be  taken 
as  a  fair  average,  especially  for  a  new  type  that  is  not  as 
refined  in  detail  as  the  H-3. 


278  FUSELAGE  DETAILS 


SIZE  OF  LONGERONS 

The  size  of  the  longerons,  that  is,  the  section,  is 
influenced  by  many  factors.  As  these  members  must 
resist  flying  loads,  the  leverage  of  elevator  flaps,  stresses 
due  to  control  wires,  landing  stresses  and  the  v^eight  of 
the  motor  and  personnel  it  is  always  advisable  to  itemize 
the  loading  and  then  prepare  a  diagram  to  obtain  the 
stresses  in  the  different  members.  This  latter  method 
is  a  method  for  a  trained  engineer,  but  an  exhaustive  de- 
scription of  the  method  of  procedure  will  be  found  in 
books  on  the  subject  of  "Strength  of  Materials."  For  the 
practical  man,  I  give  the  following  list  of  longeron  dimen- 
sions so  that  he  will  have  at  least  a  guide  in  the  selection 
of  his  material. 

The  length  of  the  fuselage  and  power  of  motor  are 
given  so  that  the  reader  can  obtain  sizes  by  comparison, 
although  this  is  a  crude  and  inaccurate  method.  As  the 
longerons  taper  from  front  to  back,  the  sizes  of  the  section 
are  given  at  the  motor  end,  and  also  at  the  tail.  The  size 
of  the  front  members  depends  principally  upon  the  weight 
of  the  motor  and  the  passenger  load,  while  the  rear 
longerons  carry  the  elevator  loads  and  the  tail  skid  shock. 
If  the  rudder  is  high  above  the  fuselage  it  introduces  a 
twisting  movement  that  may  be  of  considerable  impor- 
tance. The  loads  on  the  stabilizer,  elevators  and  the 
vertical  rudder  are  very  severe  when  straightening  out 
after  a  steep  dive  or  in  looping,  and  the  pull  on  the  con- 
trol wires  exerted  by  the  aviator  at  this  time  greatly  adds 
to  the  total  stress.  In  the  front  of  the  fuselage,  the  motor 
exerts  a  steady  torque  (twist)  in  addition  to  the  stress 
due  to  its  weight,  and  to  this  must  be  added  the  gyroscopic 
force  caused  by  the  propeller  when  the  machine  is  sud- 
denly changed  in  the  direction  of  flight.  The  combination 
of  these  forces  acting  at  different  times  makes  the  calcu- 
lation very  difficult. 


FUSELAGE  DETAILS 


279 


LONGERON     DIMENSIONS 


Name  of 
Machine 

Model 

Length 

of 
Fuse- 
lage 

Front 
Longerons 

Rear 
Longerons 

Power 
H.P. 

Material 

Front 
Longerons 

Material 

Rear 
Long'ons 

Curtiss 

Bipl. 
JX4  B 

24'-0' 

1.5"xl.25" 

1.25"xl.25" 

-90 

Ash 

Spruce 

Bleriot 

Monpl. 
"XI" 

1.13"xl.l3" 

0.88"x0.88" 

50 

Ash- 
Spruce 

Ash- 
Spruce 

Rumpler 

Bipl. 
1916 

21' -9" 

1.60"xl.20" 

0.80"x0.80" 

140 

Ash 

Spruce 

Standard 

Bipl. 
H-3 

24 '-6" 

1.18-X1.18" 

0.88"x0.88" 

135 

Ash 

Curtiss 

Bipl. 
R-2 

1.50"xl.25" 

1.00"xl.00" 

160 

Ash 

Spruce 

Curtiss 

Bipl. 
R-4 

24' - 6" 

1.63"xl.25" 

1.00"xl.00" 

200 

Ash 

Sloane 

Bipl. 
H-1 

1. 50"  X  1.50" 

1.00"xl.OO" 

125 

Ash 

Spruce 

Sopwith 

Bipl. 

24' -2" 

1.13"xl.l3" 

0.88"x0.88" 

50 

Ash 

Spruce 

Chicago 
Aero  Wks. 

Bipl. 
"Star" 

1.38"xl.38" 

1.00"xl.00" 

60 

Ash 

Spruce 

Chicago        Bipl. 
Aero  WksJ    "Jr." 

14 '-0" 

1.10"xl.lO" 

0.75"x0.7S" 

30 

Ash 

Spruce 

In  the  case  of  the  Curtiss  R-4,  the  front  longerons  taper 
down  from  the  motor  L63"xL25"  to  a  point  directly  behind 
the  pilot's  seat,  the  section  at  the  latter  point  being  L25"x 
1.25".  From  this  point  the  rear  longerons  taper  down 
to  V'xl"  at  the  tail.  At  the  motor,  the  section  is  1.63"x 
1.25".  The  longitudinals  of  the  Bleriot  monoplane  are 
laminated  and  are  built  up  of  alternate  layers  of  spruce 
and  ash.  This  is  an  old  type  of  machine  and  this  prac- 
tice has  since  been  discontinued.  It  will  be  noted  that  as 
the  power  is  increased,  the  size  of  the  front  longerons  is 
generally  increased,  although  this  is  not  always  the  case 
in  speed  machines.  The  Chicago  Aero  Works'  "Star" 
fuselage  could  easily  carry  a  90  horsepower  motor, 
although  this  size  is  not  regularly  installed. 

Pusher  Type  Fuselage  (Nacelle).  Compared  with  the 
tractor  biplane  and  the  monoplane  fuselage,  the  body  of 
the  pusher  is  very  short  and  light.  The  latter  body  simply 
acts  as  a  support  for  the  motor  and  personnel  since  the 
tail  loads  are  carried  by  the  outriggers  or  tail  booms. 
The  motor  is  located  at  the  rear  end  of  the  body  and  may 


280 


FUSELAGE  DETAILS 


be  either  of  the  air  or  water-cooled  type.  The  accom- 
panying figure  shows  a  typical  pusher  type  body,  or 
"Nacelle"  as  it  is  sometimes  called. 

The  advantages  of  the  pusher  type  for  military  service 
are  obvious.  The  observer  or  gunner  can  be  placed  imme- 
diately in  the  front  where  his  vision  is  unobstructed,  and 
where  the  angle  of  fire  is  at  a  maximum. 

Twin  Motored  Fuselage.  Twin  motored  aeroplanes 
generally  have  the  power  plants  mounted  at  a  point  about 
midway  between  the  fuselage  and  tips  of  the  wings.  In 
almost  every  case,  the  power  plants  are  of  unit  construc- 


Typical  Pusher  Body   Showing  Wings,  and  Outrigger  to  Tail   Surfaces. 


tion,  that  is  to  say,  consist  of  the  motor,  radiator  and  pro- 
peller complete  on  one  support,  only  the  fuel  and  oil  tanks 
being  mounted  in  the  fuselage.  The  fuselage  of  the  twin 
may  be  similar  in  length  and  general  construction  to  that 
of  the  tractor  biplane,  or  it  may  be  a  short  "nacelle" 
similar  to  that  used  in  the  pusher  type.  In  any  case,  the 
observer  can  be  located  in  the  extreme  front  of  the  body. 
An  interesting  and  unusual  construction  is  the  body  of 
the  Caproni  Biplane  (1916).  A  center  nacelle  carries  the 
passengers,  a  pusher  screw  being  located  at  the  rear  of 
the  central  body  as  in  the  case  of  the  pusher  biplane.  On 
either  side  of  the  center  are  the  motors  driving  the  tractor 


FUSELAGE  DETAILS  281 

screws,  each  motor  being  encased  in  a  long  tractor  type 
fuselage  that  also  supports  the  tail  surfaces.  The  latter 
fuselage  serves  to  streamline  the  motors  and  takes  the 
place  of  the  usual  outrigger  construction.  There  are 
three  bodies,  two  tractor  screws,  and  one  pusher  screw. 
Somewhat  similar  in  design  is  the  famous  German  "Billy- 
Two-Tails,"  this  machine  being  equipped  with  two 
tractor  type  bodies.  A  motor  is  located  in  the  front  of 
each  body.  Each  fuselage  is  provided  with  accommoda- 
tions for  passengers,  and  is  long  enough  to  support  the 
tail  surfaces.  The  Caproni  and  the  German  machine  are 
both  very  large  machine  and  heavily  powered. 

U.  S.  A.  Sea-Plane  Specifications  (1916).  These  gov- 
ernment specifications  cover  a  twin  motored  sea-plane 
with  a  central  nacelle.  The  body  is  arranged  so  that 
the  foreward  man  (observer)  can  operate  the  forward 
machine  gun  through  a  horizontal  arc  of  at  least  150°, 
and  through  a  vertical  arc  of  at  least  270°,  with  the  gun 
at  an  angle  of  about  75°  with  the  center  line  of  the  body. 
The  muzzle  must  be  forward  of  the  propeller  plane.  The 
rear  man  (pilot)  operates  a  machine  gun  through  a 
vertical  arc  of  at  least  150°  to  the  rear,  and  through  a 
vertical  arc  of  at  least  180°,  with  the  gun  at  an  angle  of 
about  105°  with  the  fuselage  center  line.  The  muzzle 
must  be  to  the  rear  of  the  plane  of  propeller  rotation. 

The  number  of  stays  and  other  important  connections 
which  extend  across  the  plane  of  propeller  rotation  shall 
be  reduced  to  a  minimum.  It  is  considered  advisable  to 
incorporate  in  the  design  of  the  body  such  a  structure  (in 
the  plane  and  8  inches  forward  of  propeller  plane)  as  will 
prevent  a  broken  propeller  blade  from  severing  the  main 
body.  The  system  used  in  the  construction  of  the  cage 
masts  used  on  battleships  is  suggested,  with  a  number  of 
spruce  compression  members  in  place  of  stay-wires.  The 
clearance  of  the  propeller  tips  from  the  sides  of  the  central 
body  shall  be  from  5  to  12  inches.    No  part  of  the  gas  tanks 


282  FUSELAGE  DETAILS 

shall  lie  in  the  plane  of  propeller  rotation,  nor  within  a 
space  6  inches  ahead  of  this  plane. 

A  space  extending  at  least  9  inches  back  from  the  rear 
of  the  observer's  seat,  and  entirely  across  the  body,  must 
be  left  open  and  unoccupied  in  order  that  any  desired 
instruments  can  be  installed  therein.  In  the  center  line  of 
the  body,  a  circular  hole  9  inches  in  diameter  shall  be  cut 
in  the  floor  of  the  observer's  cock-pit,  the  rear  of  the  hole 
being  5  inches  forward  of  the  forward  edge  of  the 
observer's  seat.  The  flooring  of  the  pilot's  and  observer's 
cockpits  shall  consist  of  spruce  strips  ^"x^"  spaced  at 
Yz"  intervals  along  the  longerons.  No  flooring  is  to  be 
placed  under  the  seats. 

The  safety  factor  of  the  body  and  tail  structure  shall 
not  be  less  than  2.5,  the  air  speed  being  taken  at  100  miles 
per  hour  with  the  elevator  at  an  angle  of  20°  and  the  fixed 
stabilizer  surface  at  6°.  All  wire  tension  members  not 
readily  accessible  for  inspection  and  adjustment  are  to  be 
single  strand  high  tensile  steel  wire.  All  tension  stays 
that  are  easily  accessible  shall  be  of  non-flexible  stranded 
steel  cable.  For  turnbuckle  safetying  No.  20  semi-hard 
copper  wire  shall  be  used.  All  cable  shall  be  well  stretched 
before  making  up  the  connections.  A  load  equal  to  20  or 
30  per  cent  of  the  breaking  load  shall  be  applied  for  a 
period  of  from  two  to  three  hours.  The  hard  wire 
must  undergo  a  bending  test  by  bending  at  a  right  angle 
turn  over  a  radius  equal  to  the  diameter  of  the  wire,  back 
and  forth  four  times  each  way.  No  more  than  four  sizes 
of  turnbuckles  shall  be  used  on  the  entire  aeroplane  struc- 
ture. The  strengths  and  size  numbers  of  the  turnbuckles 
will  be  as  follows :  No.  1  ^  8,000  lbs.  No.  2  =  4,600  lbs. 
No.  3  =  2,100  lbs.  No.  4=1,100  lbs.  Controls  and 
fittings  in  the  vicinity  of  the  compasses  shall,  as  much  as 
possible,  be  of  non-magnetic  material.  All  steel  plate  and 
forged  fittings  shall  be  protected  against  the  action  of 
salt  water  by  baking  enamel,  the  best  standard  three  coat 


FUSELAGE  DETAILS  283 

process  being  used.  All  covered  wiring  and  turnbuckles 
shall  be  coated  by  at  least  two  coats  of  Flexible  Com-^ 
pound. 

All  steel  tubing  shall  be  thoroughly  cleaned,  slushed 
with  mineral  oil  inside,  and  plugged  at  both  ends  by  wood 
plugs  impregnated  with  mineral  oil  or  paraffine.  All  steel 
nuts,  bolts,  pins  and  cotter  pins  shall  be  protected  by 
heavy  nickel  plating  over  copper.  All  wood  members, 
especially  faying  surfaces,  end  grain  butts,  scarfs  and 
joints,  shall  be  protected  against  the  access  of  moisture 
before  final  assembly  by  the  best  grade  of  varnish,  or  by 
impregnation  by  paraffine.  All  wood  shall  be  straight 
grained,  well  seasoned,  of  uniform  weight,  and  free  of 
knots,  pitch  pockets,  checks  or  cracks.  Spruce  to  be  of 
the  very  highest  grade  of  selected  straight,  even  grained, 
clear  spruce.  It  shall  be  air  seasoned,  preferably  for  two 
years.    Kiln  dried  wood  is  not  acceptable. 

It  is  highly  desirable  to  have  all  bolts,  pins,  plate  fittings- 
and  turnbuckle  ends  made  of  chrome  vanadium  steel 
(S.  A.  E.  Specification  6130),  heat  treated  to  obtain  the 
best  physical  characteristics.  All  parts  and  fittings  that 
must  be  bent  shall  be  heat  treated  after  all  bending  oper- 
ations are  completed,  and  by  such  a  sequence  of  treatment 
as  will  produce  the  desired  grain  and  toughness,  and 
relieve  all  stresses  due  to  the  bending.  This  includes 
sheet  and  forged  steel  fittings,  turnbuckle  ends  and  bolts 
and  pins.  All  steel  parts  and  fittings  submitted  to  stress 
or  vibration  shall  be  heat  treated  in  such  a  manner  as  to 
produce  the  highest  possible  refinement  of  grain  and  give 
the  greatest  possible  resistance  to  alternating  and  vibra- 
tory stresses.  Where  plate  fittings  are  in  contact  with 
wooden  members,  sharp  edges  next  to  the  wood  shall  be 
removed.  In  making  up  and  connecting  steel  fittings^ 
welding  shall  be  used  wherever  possible.  If  impracticable 
to  weld,  and  in  such  cases  only,  brazing  will  be  used, 
proper  heat  treatment  to  be  employed  to  restore  strength 


281  FUSELAGE  DETAILS 

and  toughness  of  metal  after  such  welding  or  brazing. 
Extreme  care  should  be  taken  to  avoid  nicking  or  kinking 
any  wire,  cable  or  fitting.  Fittings,  sheet  or  forged,  must 
be  free  from  sharp  corners  and  supplied  with  generous 
fillets. 

In  general  the  S.  A.  E.  Standards  will  be  acceptable, 
and  these  standards  for  screw  threads  shall  be  used  wher- 
ever possible.  U.  S.  Standard  threads  will  be  accepted 
where  threaded  into  cast  iron,  cast  aluminum  or  copper 
alloys.  All  nuts  and  pins  must  be  provided  with  one  or 
more  positive  and  durable  safety  devices.  In  general, 
where  it  must  be  expected  that  a  structural  fitting  will 
be  disassembled  a  number  of  times  during  the  life  of  the 
aeroplane,  castellated  nuts  with  split  pins,  in  accordance 
with  S.  A.  E.  Standards,  shall  be  used.  Wherever  this 
is  not  the  case,  pins  or  bolts  shall  be  riveted  in  a  work- 
manlike manner. 

Seats  shall  be  securely  braced  against  both  horizontal 
and  vertical  stresses.  Arrangement  and  dimensions  of 
-cock-pits  shall  be  as  nearly  as  practicable  to  that  indicated 
by  the  drawings  (not  published  in  this  chapter).  In 
addition,  if  practicable,  the  pilot  should  be  provided  with 
■quick  release  arm  rests.  Sections  of  best  grade  of  khaki 
■on  each  side  of  seats,  in  which  pockets  are  made,  should 
be  fastened  to  longerons  and  vertical  posts  in  such  a  way 
as  to  be  securely  in  place  and  yet  readily  detachable  for 
inspection  of  structural  wiring  and  fittings.  Safety  belts 
shall  be  provided  for  both  seats  and  securely  fastened. 
The  belts  shall  safely  support  at  any  point  a  load  of  2,000 
pounds  applied  as  in  practice.  Rubber  shock  absorbers 
in  the  safety  belt  system  are  considered  to  be  an  advan- 
tage. The  quick  release  device  shall  be  ^s  indicated 
in  drawings  and  shall  reliably  and  quickly  function. 
Seat  pads  shall  be  quickly  detachable  in  order  that 
they  may  be  used  as  life  preservers.  They  will  be 
filled  with  Kapok  or  other  similar  material  and  covered 


FUSELAGE  DETAILS  285 

with  real  leather  to  protect  it  against  the  action  of  salt 
water. 

Suitable  covers  shall  be  provided  over  the  top  of  the 
rear  end  of  the  fuselage.  These  must  be  easily  removed 
and  capable  of  being  securely  fastened  in  place  during 
flight.  Space  shall  be  allowed  in  the  body  directly  in  the 
rear  of  the  observer's  seat  for  the  stowage  of  the  sea 
anchor.  When  in  use,  the  sea  anchor  shall  be  attached 
by  suitable  and  convenient  fastening  hooks  to  the  two 
points  along  the  lower  longerons,  and  at  the  junction  of 
the  two  vertical  struts  in  the  rear  of  the  front  seat.  The 
structure  must  be  such  that  it  will  successfully  withstand 
the  stresses  imposed  by  the  sea  anchor.  Controls  shall 
be  of  the  standard  Deperdussin  type,  installed  in  the  rear 
cock-pit  only.  The  tanks  for  the  main  supply  of  gasoline 
shall  be  in  the  fuselage  and  located  so  that  the  longi- 
tudinal balance  will  not  be  disturbed  by  the  emptying 
of  the  tank  during  flight. 

The  above  data  is  not  in  the  exact  form  of  the  original 
specifications  and  is  not  complete,  but  gives  only  the 
specifications  that  affect  the  design  of  the  body.  These 
were  picked  out  part  by  part  from  the  original. 

Army  Specification  1003  (Speed  Scout).  These  speci- 
fications cover  the  design  of  land  machines,  the  extracts 
given  here  referring  only  to  the  safety  factor.  Body  for- 
ward of  the  cockpit  shall  be  designed  for  safety  factor 
of  10  over  static  conditions,  with  the  propeller  axis  hori- 
zontal. Body  in  rear  of  cockpit  shall  be  designed  to  fail 
under  loads  not  less  than  those  imposed  under  the  fol- 
lowing conditions  : 

(a)  Dynamic  loading  of  5  as  the  result  of  quick  turns 
in  pulling  out  of  a  dive,  (b)  Superposed  on  the  above 
dynamic  loading  shall  be  the  load  which  it  is  possible  to 
impose  upon  the  elevators,  computed  by  the  following 
formula  :  L  =  0.005 AV-,  where  A  is  the  total  area  of  the 
stabilizing  surface   (elevators  and  fixed  surface),  and  V 


286  FUSELAGE  DETAILS 

is  the  horizontal  high  speed  of  the  machine.  The  units 
are  all  in  the  metric  system,  (c)  Superposed  on  this 
loading  shall  be  the  force  in  the  control  cables  producing 
compression  in  the  longerons. 

Fuselage  Covering.  Disregarding  the  monocoque  and 
veneer  constructed  types  of  fuselage,  the  most  common 
method  of  covering  consists  of  a  metal  shell  in  the  for- 
w^ard  end,  and  a  doped  linen  covering  for  that  portion 
of  the  body  that  lies  to  the  rear  of  the  rear  seat.  The 
metal  sheathing,  which  may  be  of  sheet  steel  or  sheet 
aluminum,  generally  runs  from  the  extreme  front  end  to 
the  rear  of  the  pilot's  cockpit.  Sheet  steel  is  more  com- 
mon than  aluminum  because  of  its  stiffness.  Military 
machines  are  usually  protected  in  the  forward  portions 
of  the  fuselage  by  a  thin  armor  plate  of  about  3  milli- 
meters in  thickness.  This  is  a  protection  against  rifle  bul- 
lets and  shrapnel  fragments,  but  is  of  little  avail  against 
the  heavier  projectiles.  Armor  is  nearly  always  omitted 
on  speed  scouts  because  of  its  weight.  Bombers  of  the 
Handley-Page  type  are  very  heavily  plated  and  this  shell 
can  resist  quite  large  calibers. 

The  fabric  used  on  the  rear  portion  of  the  fuselage  is 
of  linen  similar  to  the  wing  covering,  and  like  the  wing 
fabric  is  well  doped  with  some  cellulose  compound  to 
resist  moisture  and  to  produce  shrinkage  and  tautness. 
On  the  sides  and  bottom  the  fabric  is  supported  by  very 
thin,  light  stringers  attached  to  the  fuselage  struts.  On 
the  top,  the  face  is  generally  curved  by  supporting  a 
number  of  closely  spaced  stringers  on  curved  wooden 
formers.  The  formers  are  generally  arranged  so  that 
they  can  be  easily  removed  for  the  inspection  of  the 
wire  stay  connections  and  the  control  leads.  On  some 
machines  the  top  of  the  fuselage  consists  entirely  of  sheet 
metal  suported  on  formers,  while  in  others  the  metal  top 
only  extends  from  the  motor  to  the  rear  of  the  rear  cock- 
pit. 


CHAPTER  XIII. 
CHASSIS   CONSTRUCTION. 

General  Notes.  The  chassis  or  landing  gear  carries  the 
weight  of  the  aeroplane  when  resting  on  or  running  over 
the  ground,  and  is  subjected  to  very  heavy  shocks,  espe- 
cially when  landing.  It  is  provided  with  pneumatic  tired 
wheels,  an  elastic  shock  absorbing  device,  and  the  struc- 
tural members  that  connect  the  axle  with  the  fuselage. 
In  some  forms  of  landing  gear,  the  wheels  are  supple- 
mented by  long  horizontal  skids  which  serve  to  support 
the  machine  after  the  elastic  shock  absorbers  are  fully  ex- 
tended or  when  the  wheels  collapse.  The  skids  also  pro- 
tect the  aeroplane  in  cases  where  the  wheels  run  into  a 
ditch  and  also  prevent  the  machine  from  nosing  over  in  a 
bad  landing.  Since  the  skids  and  their  structural  members 
cause  a  high  resistance  they  are  now  seldom  used  except 
on  the  larger  and  slower  machines.  In  running  over  the 
ground,  or  in  making  a  hard  landing,  part  of  the  shock  is 
taken  up  by  the  deflection  of  the  tires  and  part  by  the 
deflection  of  the  shock  absorber.  The  greater  the  move- 
ment of  the  tires  and  absorber,  the  less  will  be  the  stress 
in  the  frame. 

In  the  majority  of  cases,  the  shock  absorbers  consist  of 
rubber  bands  or  cords,  these  being  wound  over  the  axle 
and  under  a  stationary  part  of  the  chassis  members.  Since 
rubber  is  capable  of  absorbing  and  dissipating  a  greater 
amount  of  energy  per  pound  of  weight  than  steel,  it  is  the 
most  commonly  used  material.  Rubber  causes  much  less 
rebound  or  **kick"  than  steel  springs.  The  principal  ob- 
jection to  rubber  is  its  rotting  under  the  influence  of  sun- 
light, or  when  in  contact  with  lubricating  oil.    The  move- 

287 


288 


CHASSIS  CONSTRUCTION 


ment  of  the  axle  tube  is  generally  constrained  by  a  slotted 
guide  or  by  a  short  radius  rod. 

The  design  of  a  suitable  chassis  is  quite  a  complicated 
problem,  for  the  stresses  are  severe,  and  yet  the  weight 
and  resistance  must  be  kept  at  a  minimum.  In  running 
over  rough  or  soft  ground  for  the  ''Get  off,"  the  shocks 


Fig. 


1.     "V"  Type  Chassis  as  Applied  to  "Zens' 
•'Flight." 


Monoplane.     Courtesy 


and  vibration  must  be  absorbed  without  excessive  stress 
in  the  framework,  and  without  disturbing  the  balance  or 
poise  of  the  machine.  There  must  be  little  tendency  to- 
ward nosing  over,  and  the  machine  must  be  balanced 
about  the  tread  so  that  side  gusts  have  little  tendency  in 
throwing  the  machine  out  of  its  path.  It  must  be  simple 
and  easily  repaired,  and  the  wheels  must  be  large  enough 
to  roll  easily  over  moderately  rough  ground. 


CHASSIS  CONSTRUCTION 


289 


Types  of  Chassis.  The  simplest  and  most  extensively 
used  landing  gear  is  the  "Vee"  type  shown  by  Fig.  1,  and 
is  equally  applicable  to  monoplanes,  biplanes  or  triplanes. 
Primarily,  the  Vee  chassis  consists  of  two  wheels,  an 
axle,  a  rubber  shock  absorber,  and  two  sets  of  Vee  form 
struts.  The  chassis  shown  by  Fig.  1-a  is  that  of  the  Hansa- 
Brandenburg  and  is  typical  of  biplane  chassis.  The  wind- 
ing of  the  rubber  cord  and  the  arrangement  of  the  chassis 


itirren.ngBo4.  3  5  mm 
Kimntboa 

lOrrjin  t>o(t 
t-Smep  ducfftt 


,  tlmm.  Striper  ai^d  Cable 
,  IQmn.  bcnii 


Shock  Absorber  Shock  Absorber- 

Fig.  1-a.     "V"  Type  Chassis  Used  on  Hansa-Brandenburg  Biplane. 

Struts  are  clearly  shown.  The  two  struts  are  connected  at 
the  bottom  by  a  metal  fitting,  and  the  rubber  is  wound 
over  the  axle  and  under  this  fitting.  No  guiding  device 
is  used  for  the  axle,  the  machine  being  freely  suspended 
by  the  chord.  The  struts  are  made  as  nearly  streamline 
form  as  possible. 

Fig.  2  is  a  front  view  of  a  typical  Vee  chassis,  and  Fig.  3 
is  side  view  of  the  same  device,  the  same  reference  letters 
being  used  in  each  view.  The  vertical  struts  C  run  from 
the  fuselage  at  F  to  the  connecting  axle  guide  plate  G. 
The  wheels  W-W  are  connected  with  the  steel  tube  axle 


290 


CHASSIS  CONSTRUCTION 


A,  and  the  struts  are  braced  against  side  thrust  by  the 
cross-tube  D  and  the  stay  wire  braces  B-B.  In  Fig.  3 
the  metal  fitting  G  is  provided  with  the  guiding  slot  S  for 
the  axle  A.  The  elastic  rubber  cord  absorber  passes  over 
the  axle  and  is  fastened  to  the  plate  G  by  the  studs  I. 
Fig.  4  is  a  side  view  of  the  chassis  of  the  Lawson  trainer, 
which  like  many  other  primary  training  machines,  uses 
a  front  pilot  wheel  to  guard  against  nosing  over.  The 
rear  two  wheels  (W)  are  elastically  supported  between 


rxs.S 


rr&-2. 


Figs.  2-3.     Typical    "V"    Chassis    With    Axle    Guide. 

the  Vee  struts  C  and  F,  while  the  front  wheel  X  is  at- 
tached to  the  fuselage  by  the  vertical  strut  E,  and  to  the 
rear  wheel  frame  by  the  tube  G.  It  will  be  noted  that  the 
front  wheel  is  smaller  than  the  rear  main  wheels,  as  this 
wheel  carries  but  little  load.  The  tail  skid  T  is  hinged 
to  the  fuselage  and  is  provided  with  elastic  cord  at  the 
upper  end  so  that  the  shock  is  reduced  when  the  tail 
strikes  the  ground.  Fig.  5  shown  directly  above  the  Law- 
son  trainer,  is  the  complete  assembly  of  the  Hansa-Bran- 
denburg  already  described.  The  tail  skid  of  the  Hansa- 
Brandenburg  is  indicated  by  T. 

The  metal  shod  ash  skid  stick  is  hinged  to  the  lower 
face  of  the  fuselage,  and  at  the  upper  end  is  attached  to  a 
stationary  fuselage  member  through  four  turns  of  elastic 


CHASSIS  CONSTRUCTION 


291 


cord.  When  the  skid  strikes  an  obstacle  the  rubber  gives 
and  allows  the  tail  to  move  in  relation  to  the  ground.  By 
this  arrangement  the  greater  part  of  the  device  is  enclosed 
within  the  fuselage  and,  hence,  produces  little  head  resist- 
ance. 

Fig.  7  is  the  skid  chassis  of  the  Farman  biplane  which 


Fig    4 


(Below).      Lawson    Training    Tractor    Biplane.     Fig.    S     (Above). 
ansa-Brandenburg  Fighting  Biplane  Showing  Chassis  and  Tail  Skid  (t). 


shows  clearly  the  arrangement  of  the  skids  and  the  shock 
absorbing  suspension.  A  metal  bridge  is  attached  to  the 
axle,  and  a  series  of  short  rubber  bands  are  used  in  con- 
necting the  axle  bridge,  and  the  bridge  on  the  skid.  A 
triangular  tubular  radius  rod  is  attached  to  the  axle  and 
hinged  to  the  skid.  This  restrains  the  travel  of  the  axle 
in  a  fore  and  aft  direction.  Another  form  of  skid  shock 
absorber  is  given  by  Fig.  8,  in  which  the  rubber  rings 
pass  over  a  spool   on  the  axle.     The  guiding  links  or 


292  CHASSIS  CONSTRUCTION 

radius  rods  on  the  inside  of  the  skids  regulate  the  axle 
travel.  In  general,  the  use  of  a  radius  rod  is  not  de- 
sirable as  it  transmits  a  percentage  of  the  shock  to  the 
machine. 


Fig.   7.      (Left).  _   Farman  Skid  Type  rba';^.:?.     Fig.  S.     Another  Type  of  Skid 
Chassis  in  Which  the  Axle  Is  Guided  by  a  Radius  Rod  or  Lever. 


Fig.  9.  Chassis  Details  of  the  Xieuport  Monoplane.  This  Has  a  Central 
Skid  and  Uses  an  Automobile  Type  Steel  Spring  Instead  of  Rub- 
ber Cord.     Fig.   10  Is  a  Detail  of  the  Nieuport  Spring.      (At  Right.) 

Fig.  9  is  an  older  form  of  Nieuport  monoplane  chassis, 
a  steel  cross  spring  being  used  in  place  of  the  usual  rubber 
bands.  This  is  simple,  but  comparatively  heavy,  and  is 
subject  to  frequent  spring  breakage.  To  guard  against 
spring  failure,  a  long  ash  skid  is  placed  under  the  axle. 
The  spring  system  is  connected  with  the  body  by  three 
sets  of  oval  steel  struts.  An  old  type  of  Curtiss  chassis  is 
given  by  Fig.  11.  This  has  been  widely  used  by  amateurs 
and  exhibition  flyers,  but  requires  a  fairly  smooth  landing 
ground  as  there  are  no  shock  absorbers.  The  only  shock 
absorption  is  that  due  to  the  deflection  of  the  tires.    The 


CHASSIS  CONSTRUCTION 


293 


extreme  forward  position  of  the  front  wheel  effectually 
prevents  any  tendency  toward  nosing  over  when  landing. 
A  Standard  H-3  shock  absorbing  system  is  given  by 
Fig.  12.  This  has  a  bracket  or  hanger  attached  to  the 
axle  over  which  the  elastic  cord  i?  wrapped.    The  cord  is 


Fig.    11.     An   Old  Type  of  Curtiss   Exhibition   Chassis  With   Three  Wheels. 


Fig.  12  (Left).     Standard  H-3  Shock  Absorber. 
Cord  on  Axle. 


(Right).     Rubber 


wrapped  in  continuous  turns  between  the  axle  hanger 
and  the  bottom  of  the  Vee  support  members.  As  shown, 
the  upper  streamlined  bar  is  the  axle,  while  the  lower  is 
the  cross  bar  brace  which  serves  to  hold  the  lower  ends 
of  the  U's.  I  am  indebted  to  "Aerial  Age"  for  this  cut. 
In  order  to  guide  the  axle  in  a  straight  line  in  its  up  and 


294 


CHASSIS  CONSTRUCTION 


down  movement,  two  radius  links  are  attached  between 
the  axle  and  the  front  vertical  strut.  One  decided  ad- 
vantage of  the  "Standard  construction  is  that  the  cords 
are  wound  without  crossing  the  strands,  thus  reducing 
cutting  and  wear  between  the  cord  turns. 

Fig.  13  is  a  variation  of  Fig.  12,  the  cord  being  wound 
directly  around  spools  on  the  axle  and  the  lower  station- 
ary cross  tube.  Th  axle  is  guided  by  a  slot  in  the  guide 
plate  at  the  right,  while  end  motion  is  controlled  by  a 
radius  link.    Fig.  14  is  the  double  wheel  arrangement  of  a 


Fig.  14.     Chassis  for  Twin  Motored  Biplane  of  Bombing  Type. 

large  "Twin"  bombing  plane.  Two  wheels  are  placed 
directly  under  each  of  the  motor  units  so  that  a  portion  of 
the  load  is  communicated  to  the  chassis  by  tubes.  Diag- 
onal tubes  transmit  the  body  load  to  the  chassis. 

Folding  Chassis.  Owing  to  the  .great  relative  resist- 
ance of  the  chassis  it  has  been  suggested  by  many  de- 
signers to  provide  a  folding  frame  which  will  atuomatic- 
ally  fold  up  into  the  body  after  the  machine  has  left  the 
ground.  This  would  be  a  decided  advantage  but  the  gear 
is  complicated  and  probably  not  altogether  reliable. 

Height  of  Chassis.  The  height  of  the  chassis  is  made 
as  small  as  possible  with  a  sufficient  clearance  for  the 
propeller  tips.  It  is  usual  to  have  the  tips  of  the  propeller 
blades  clear  the  ground  by  from  10  to  12  inches  when  the 
aeroplane  is  standing  with  the  body  in  a  horizontal  posi- 
tion.   Any  smaller  clearance  is  almost  certain  to  result  in 


CHASSIS  CONSTRUCTION 


295 


broken  blades  when  landing  at  a  sharp  angle  or  when  run- 
ning through  high  grass.  If  the  chassis  is  excessively 
high  the  resistance  will  be  high  and  the  machine  is  also 
likely  to  be  top  heavy. 

Location  of  Wheels.  The  exact  location  of  the  wheels, 
in  a  fore  and  aft  direction,  is  of  the  greatest  importance. 
If  they  are  too  far  ahead  of  the  center  of  gravity,  too  much 


Figs.   15-16.     Methods  of  Calculating  Wheel  Position  on  Two   Wheel   Chas- 
sis.    This  Is  an  Important  Item  in  the  Design  of  an  Aeroplane. 

w^eight  will  be  placed  on  the  tail  skid  and  excessive  run- 
ning will  be  required  to  get  the  tail  ofif  the  ground.  If 
the  wheels  are  too  far  back,  the  machine  will  be  likely  to 
nose  over  when  landing  or  running  over  the  ground.  In 
any  case,  the  wheels  must  be  w^ell  ahead  of  the  center  of 
gravity  so  that  the  weight  will  resist  a  forward  overturn- 
ing moment.  In  the  majority  of  orthogonal  biplanes,  in 
which  the  leading  edges  of  the  upper  and  lower  wings  are 
on  the  same  vertical  line,  the  center  of  the  wheel  is  from 
3  to  6  inches  back  of  the  leading  edges.     In  staggered 


296  CHASSIS  CONSTRUCTION 

biplanes  the  wheel  center  is  from  6  inches  to  one  foot  in 
front  of  the  lower  leading  edge.  This  difference  is  caused 
by  the  fact  that  the  center  of  gravity  is  nearer  the  leading 
edge  of  a  staggered  wing  than  with  the  orthogonal  type, 
and  hence  the  wheels  must  be  further  forward. 

Fig  15  (upper  diagram)  shows  the  conditions  when  the 
machine  is  running  over  the  ground  with  the  body  hori- 
zontal. The  vertical  line  a-a  passing  through  the  center 
of  gravity  C  G  is  a  distance  N  from  the  center  of  the 
wheel.  The  weight  acting  down  has  a  tendency  to  pull 
the  tail  down,  this  moment  being  equal  to  the  weight  of 
the  machine  multiplied  by  the  distance  N,  or  W  x  N.  The 
elevator  flap  M  exerts  a  lifting  force  K,  which  acts 
through  the  lever  arm  L,  and  opposes  the  moment  due  to 
the  weight.  The  force  K  must  be  equal  to  K  =  WX/L. 
The  distance  I  is  the  distance  of  the  wheel  center  line  from 
the  entering  edge  of  the  wing.  The  weight  on  the  tail 
skid  S  when  the  machine  is  resting  on  the  ground  will  be 
equal  to  S  =  WN/M,  and  this  may  range  anywhere  from 
40  to  200  pounds,  according  to  the  size  of  the  aeroplane. 

Fig.  16  illustrates  a  principle  of  wheel  location  ad- 
vanced by  Capt.  Byron  Q.  Jones,  and  published  in  "Avia- 
tion and  Aeronautical  Engineers,"  Nov.  16,  1916.  The 
body  is  shown  in  a  horizontal  position  with  the  propeller 
axis  X-X  horizontal.  The  center  of  gravity  is  at  G  on 
X-X,  the  weight  acting  down  as  at  P  with  the  line  pro- 
longed meeting  the  ground  line  at  B.  A  line  E-E  is  a 
line  drawn  tangent  to  the  wdieels  and  the  tail  skid  at  D, 
the  angle  of  this  line  with  the  ground  determining  the 
maximum  angle  of  incidence.  E-E  is  the  ground  line 
when  the  machine  is  at  rest.  For  the  best  conditions, 
Capt.  Jones  finds  that  the  line  connecting  the  point  of 
tagency  C,  and  the  center  of  gravity  at  G,  should  make  an 
angle  of  13  degrees  and  10  minutes  with  the  vertical  GB 
dropped  through  the  center  of  gravity.  With  the  line  GA 
drawn  perpendicular  to  the  resting  line  E-E,  the  angle 


CHASSIS  CONSTRUCTION  297 

BGA  should  be  10  degrees  as  nearly  as  possible.  This  is 
for  a  two-wheel  \'ee  chassis,  but  with  a  third  front  wheel 
as  with  the  training  of  type  the  angle  CGB  can  be  made 
less.  Capt.  Jones  has  found  that  with  the  wheels  in  the 
above  location  there  will  be  no  tendency  to  nose  over  even 
with  very  poor  landings,  and  this  method  has  been  applied 
to  the  training  machines  at  the  San  Diego  Signal  Corps 
aviation  school.  If  the  angle  BGA  is  greater  than  10 
degrees  it  is  difficult  to  "taxi"  the  machine  on  the  ground, 
this  tending  to  make  the  machine  spin  or  turn  into  the 
wind.  Capt.  Jones  claims  that  a  two-wheel  chassis  ar- 
ranged according  to  these  rules  is  superior  to  the  three- 
wheel  type  for  training  purposes  since  the  tendency 
toward  spinning  is  less. 

The  location  of  the  tail  skid  S  should  be  such  that  the 
elevator  and  rudder  surfaces  are  well  off  the  ground  with 
the  skid  fully  deflected,  and  yet  the  skids  must  be  low 
enough  to  permit  of  the  maximum  angle  of  incidence  or 
an  angle  of  EXX=  10  degrees.  To  a  certain  extent,  the 
maximum  angle  of  incidence  determines  the  chassis 
height.  If  the  angle  EXX  is  made  greater  than  the  great- 
est angle  of  incidence,  the  wings  can  be  used  as  air  brakes 
in  bringing  the  machine  to  a  quick  stop  after  landing. 

The  track,  or  the  distance  between  the  centers  of  the 
wheels  measured  along  the  axle,  must  be  about  1/7  or  0.15 
of  the  span  of  the  lower  w^ng.  This  makes  the  track  vary 
from  5  to  7  feet  on  the  usual  types,  and  as  high  as  15  feet 
on  the  large  bombing  planes.  The  track  must  be  great 
enough  to  prevent  overturning  when  making  a  landing  on 
soft  ground  or  with  a  cross  wind.  If  the  track  is  excessive, 
there  will  be  a  heavy  spinning  moment  in  cases  where  one 
wheel  strikes  a  depression  or  soft  spot  in  the  ground. 

Shock  Absorbers.  The  axle  movement  allowed  by  the 
elastic  shock  absorbers  and  guiding  appliances  averages 
from  5  to  6  inches.  The  greater  the  movement,  the  less 
will  be  the  stresses  induced  by  a  given  drop,  but  in  prac- 


298  CHASSIS  CONSTRUCTION 

tice  tlie  movement  is  generally  limited  by  considerations 
of  chassis  height  and  propeller  clearance.  It  can  be 
proved  that  a  movement  of  5  inches  will  produce  a 
maximum  stress  equal  to  8.6  times  the  weight  of  the 
machine  under  conditions  of  a  one-foot  drop,  while  with 
an  absorber  movement  of  6  inches  the  stress  is  reduced 
to  7.5  times  the  weight.  This  calculation  takes  the  tire 
deflection  into  consideration.  With  the  absorber  move- 
ment limited  to  one  inch,  the  stress  may  be  as  high  as 
35  times  the  weight  of  the  machine. 

F  ^  W  (2  -f  2.77/x)  where  W  =  weight  of  machine  in 
pounds,  F:^the  stress  produced  by  the  fall,  and  x  =  the 
absorber  movement  in  inches. 

Landing  Gear  Wheels.  The  wheels  are  generally  of 
the  tangent  laced  ware  spoke  type,  and  are  enclosed  with 
discs  to  reduce  the  resistance.  They  must  have  very  wide 
hubs  to  resist  the  heavy  end  stresses  caused  by  landing 
sidewise.  The  length  of  the  hub  should  be  at  least  twice 
the  diameter  of  the  tire  and  a  greater  width,  say  three 
times  the  tire  diameter,  is  preferable.  The  narrow  hubs 
used  on  motorcycle  wheels  are  not  safe  against  side  blows, 
although  they  may  be  capable  of  withstanding  the  vertical 
load.  The  wheels  are  rated  according  to  the  outside 
diameter  over  the  tire,  and  by  the  diameter  of  the  tire 
casing.  A  26"  x  4"  wheel  signifies  that  the  outside 
diameter  is  26  inches  with  a  casing  diameter  of  4  inches. 


Wheel 
Diam. 

Casing 
Diam. 

Ca 

Trying  Capacity 
Per  Casing 

Inflation 
Pressures 

Rim-Sizes 

26 
20 
26 
26 

3 
4 
4 
5 

375  Lbs. 
750     " 
750     " 
1,225     " 

45  Lbs. 
60     •• 
60     " 
70     " 

26x3 
20x4 
26x4 
26x5 

The  26x4  tires  are  used  on  the  majority  of  training 
machines  of  the  two-wheel  type,  while  a  20x4  wheel  is 
used  for  the  front  wheel  of  the  three-wheel  trainer.  Two 
larger  sizes,  30x4  and  34x4,  have  also  been  used  to 
some  extent,  particularly  on  the  Ackerman  spring  wheels. 


CHAPTER  XIV. 
ESTIMATION  OF  WEIGHT. 

Effect  of  Weight.  Weight  is  an  all  important  consid- 
eration and  is  most  difficult  to  estimate  unless  one  has 
accurate  data  on  existing  machines  of  the  same  type. 
The  total  weight  in  flying  order  depends  upon  the  useful 
load  to  be  carried,  and  upon  the  weight  of  the  power 
plant.  The  weight  of  the  latter  varies  both  with  the  useful 
load  and  with  the  speed,  climb,  and  duration  of  flight. 
The  type  of  aeroplane  determines  the  relative  head  resist- 
ance which  again  reflects  back  to  the  weight  of  the  power 
plant. 

The  only  reason  for  the  existence  of  an  aeroplane  is  to 
carry  a  certain  useful  load  for  a  given  distance,  and  this 
useful  load  is  the  basis  of  our  weight  calculations.  The 
basic  useful  load  consists  of  the  passengers  and  cargo, 
although  in  some  specifications  the  live  load  may  be  con- 
strued as  including  the  weight  of  the  fuel,  oil  and  instru- 
ments, and  in  the  case  of  military  aeroplanes,  the  weight 
of  the  armament,  armor,  ammunition,  wireless  and 
cameras.  For  comparison,  the  elements  constituting  the 
live  load  should  always  be  specified. 

For  a  given  horsepower,  speed  and  climb,  it  is  obvious 
that  the  dead  or  structural  weight  should  be  at  a  minimum 
for  a  maximum  live  load  capacity.  The  dead  load  carried 
in  present  aeroplanes  will  be  undoubtedly  reduced  in  the 
future  by  the  adoption  of  lighter  and  stronger  materials, 
better  methods  of  bracing,  and  by  reductions  in  the  weight 
of  the  power  plant.  Just  as  the  automobile  industry  de- 
veloped light  and  powerful  materials  of  construction,  so 

299 


300  WEIGHT  ESTIMATES 

will  the  aeroplane  designer  develop  more  suitable  ma- 
terials for  the  aircraft.  While  the  present  power  plant  has 
been  refined  to  a  remarkable  extent  when  compared  with 
the  older  types,  it  is  still  far  from  the  lowest  possible  limit. 
At  present  the  complete  power  unit — the  motor,  radiator, 
propeller,  water,  etc. — wdll  weigh  from  2  to  5  pounds  per 
horsepower. 

With  a  given  aeroplane,  the  performance  is  determined 
by  the  total  w^eight  and  power.  The  duration  and  flight 
range  can  be  increased  by  increasing  the  fuel  weight  at 
the  expense  of  the  passenger  or  cargo  weight.  The  power 
available  for  climbing  is  the  excess  of  the  total  power  of 
the  motor  over  the  power  required  for  horizontal  flight. 
Since  the  power  for  horizontal  flight  depends  principally 
upon  the  weight,  it  is  at  once  evident  that  the  weight  is 
a  regulating  factor  in  the  climbing  speed.  In  fact  the 
climbing  speed  may  be  almost  directly  determined  from 
the  weight  carried  per  horsepow^er  at  normal  flight  speed. 
A  fast  climbing  scout  may  weigh  from  8  to  12  pounds  per 
horsepower,  w^hile  the  large  low  climbing  m.achine  will 
weigh  from  16  to  20  pounds  per  horsepower,  the  respective 
climbing  speeds  being  approximately  1,200  and  350  feet 
per  minute. 

Fuel  Efficiency  and  Weight.  The  efficiency  of  the 
motor,  or  its  fuel  consumption  for  a  given  output,  has  a 
very  marked  effect  upon  the  total  w^eight  of  the  aeroplane. 
Under  certain  conditions  a  very  light  motor  with  a  high 
fuel  consumption  will  often  contribute  more  to  the  total 
weight  than  a  heavier  but  more  economical  motor.  In 
short  flights,  up  to  3  hours,  the  very  light  rotating  cylinder 
motor  with  its  high  fuel  consumption  probably  gives  the 
least  total  w^eight,  but  for  longer  flights  the  more  efficient 
and  heavier  water-cooled  type  is  preferable.  For  flights 
of  over  three  hours  the  fuel  w^eight  is  a  considerable  per- 
centage of  the  total  weight.  The  proper  motor  for  any 
machine  must  be  selected  by  computing  the  w^eight  of  the 


WEIGHT  ESTIMATES  301 

fuel  and  oil  required  for  a  given  duration  and  then  adding 
this  to  the  total  weight  of  the  engine  and  its  cooling 
s\stem. 

Distribution  of  Weight.  Practically  the  only  way  to 
predict  the  weight  of  a  proposed  machine  is  to  compare 
it  with  a  similar  existing  type.  After  the  ratio  of  the 
useful  load  to  the  total  load  has  been  determined,  the  use- 
ful load  of  the  proposed  machine  can  be  divided  by  the 
ratio  factor  to  obtain  the  total  weight.  It  should  be  noted 
in  this  regard  that  if  the  proposed  machine  is  much  larger 
than  the  nearest  existing  example,  a  liberal  allowance 
must  be  made  to  compensate  for  the  increase  in  the  pro- 
portional weight  of  the  structural  members.  There  have 
been  many  mathematical  formulas  advanced  for  predict- 
ing the  weight,  but  these  are  very  inaccurate  in  the  ma- 
jority of  cases. 

As  a  rough  estimate,  based  on  a  number  of  successful 
machines,  the  weight  of  the  actual  aeroplane  structure 
without  power-plant,  live  load,  fuel,  oil,  or  tanks,  is  very 
nearly  32  per  cent  (0.32)  of  the  total  weight.  The  remain- 
ing 68  per  cent  is  divided  up  among  the  power-plant,  fuel 
and  live  load.  Thus,  the  aeroplane  structure  proper  of  a 
machine  weighing  2000  pounds  total  will  be  2000  x  0.32  = 
640  pounds.  Taking  the  weight  of  the  power  plant,  tanks 
and  piping  at  28  per  cent,  the  total  dead  load  of  the  bare 
machine  without  fuel  or  oil  will  be  60  per  cent  of  the 
total.  With  a  training  aeroplane  built  for  a  6-hour  flight, 
the  fuel  and  oil  will  approximate  16  per  cent,  so  that  the 
total  percentage  possible  for  the  crew  and  cargo  will  be 
24  per  cent.  With  a  given  live  load,  the  total  load  can 
now  be  calculated  by  dividing  the  live  load  by  its  per- 
centage. Using  the  above  value,  for  example,  the  total 
weight  in  order  of  flight  with  a  live  load  of  720  pounds 
becomes  :  W  =  720/0.24  =  3000  pounds. 

In  government  specifications  the  total  weight  of  the 
pilot  and  passenger  are  taken  at  330  pounds,  or  165  pounds 


802  WEIGHT  ESTIMATES 

per  man.  Gasoline  and  oil  are  for  a  4-hour  flight.  A  safer 
average  figure  will  be  170  pounds  per  man,  and  a  fuel 
allowance  of  6  hours.  The  floats  of  a  seaplane  or  flying 
boat  bring  the  percentage  of  the  dead  load  much  higher 
than  with  the  land  type  of  chassis. 

The  following  table  will  give  an  idea  as  to  the  weight 
distribution  expressed  both  in  pounds,  and  as  a  percentage 
of  the  total  weight.  It  covers  a  wide  range  of  types,  vary- 
ing from  the  training  types  Curtiss  JN-4B  and  the  Stand- 
ard H-3,  to  the  Handley-Page  Giant  bomber  and  the 
Nieuport  speed  scout.  The  average  values  found  by  Hun- 
saker  for  a  number  of  machines  weighing  in  the  neighbor- 
hood of  2500  pounds  is  given  in  the  fourth  column.  Under 
each  heading  are  the  actual  weights  and  the  percentages 
of  the  total  weight  for  each  item.  Items  marked  (*) 
include  both  gasoline  and  oil.  Mark  (C)  is  the  power 
plant  complete,  and  (@)  includes  radiator. 

Weight  Per  Horsepower.  As  already  explained,  the 
weight  carried  per  horsepower  varies  with  the  type  of 
machine.  When  the  total  weight  is  determined  for  any 
aeroplane,  the  power  requirements  can  be  calculated  by 
dividing  the  total  weight  by  the  weight  per  horsepower 
ratio.  A  fair  value  for  a  training  or  exhibition  machine  is 
from  18  to  20  pounds  per  horsepower,  while  for  a  very 
high  speed  machine,  such  as  a  chaser,  the  weight  will  be 
taken  at  10  pounds  per  horsepower.  For  two-seater  fight- 
ers 16  to  18  pounds  is  fair  practice.  For  a  comparison  of 
the  horsepower-weight  ratios  used  on  different  well- 
known  machines  see  tables  in  Chapter  II.  Thus,  if  our 
total  weight  is  found  to  be  2400  pounds  as  determined 
from  the  above  table,  and  if  this  is  a  training  machine,  the 
horsepower  will  be:  2400/20^120  horsepower.  Using 
the  same  total  weight,  but  powered  for  two-seater  fighter 
conditions,  the  power  will  be  increased  to  2400/16=  150 
horsepower.  As  a  scout  the  power  will  be  increased  still 
further  to  2400/10  =  240  horsepower. 


WEIGHT  ESTIMATES 


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304  WEIGHT  ESTIMATES 

As  a  problem  in  solving  the  weight  and  horsepower 
from  the  data,  we  will  assume  that  we  are  to  design  a 
two-seater  fighter  with  a  total  useful  load  of  1200  pounds. 
This  load  consists  of  the  following  items :  Personnel  (2) 
=  330  pounds ;  gas  and  oil  =  500  pounds  ;  guns  and  am- 
munition =  370  pounds.  The  nearest  example  that  we 
have  to  this  live  load  is  that  of  the  Standard  H-3,  which 
carries  744  pounds  and  in  which  the  percentage  of  live 
load  is  28.1  per  cent.  As  our  machine  will  be  somewhat 
larger,  we  will  not  be  far  from  the  truth  if  we  take  the  per- 
centage as  0.27  instead  of  0.281.  The  total  weight,  in  fly- 
ing order,  will  now  be  1200/0.27  =  4440  pounds.  At  16 
pounds  per  horsepower  the  motor  will  be  :  4440/16  =  277 
horsepower. 

An  empirical  formula  for  a  high-speed  scout  was  set 
forth  in  ''Aviation  and  Aeronautical  Engineering"  by  D. 
W.  Douglas.  This  is  based  on  the  horsepower  unit.  A 
unit  wing  loading  of  8.45  pounds  per  square  foot,  and  a 
low  speed  of  55  miles  per  hour  was  assumed.  The  wing 
section  chosen  was  the  U.  S.  A.-l.  In  the  formula,  H  = 
horsepower : 

Power  plant  weight  =  3  H. 

Chassis  weight  ^0.7  H. 

Tail  weight  =  0.25  H. 

Fuel  for  2.25  bourse  1.4  H. 

Military  load  =  250  pounds. 

Tanks  and  piping  =  0.42  H. 

Fuselage  weight  =  1.84  H. 

Wing  weight  =  1  lb.  sq.  ft. 

Propeller  =  2.8/H. 

(Total)  =  7.61  H  +  2.5/H  +  250  =  Weight  of  aero- 


7.45 
plane  fully  loaded  in  the  order  of  flight. 

Weight  of  Wings.     The  weight  of  the  wings  depends 
upon  the  span,  very  small  machines  having  wings  that 


WEIGHT  ESTIMATES  305 

weigh  only  0.38  pounds  per  square  foot,  while  the  wings 
of  very  large  machines  may  run  as  high  as  1.1  pounds  per 
square  foot.  For  average  size  biplanes  from  0.75  to  0.80 
pounds  per  square  foot  would  probably  be  safe — that  is, 
for  areas  ranging  from  450  to  550  square  feet.  The 
weight  of  the  upper  wing  of  the  Nieuport  is  0.815  pounds 
per  square  foot,  while  the  lower  wing  (short  chord)  is 
0.646  pounds  per  square  foot.  The  wings  of  the  Standard 
H-3  trainer  will  average  0.77  pounds  per  square  foot, 
the  lower  wing  and  center  section  being  heavier  than  the 
upper  wing.  The  wings  of  the  Curtiss  JN-4B  will  aver- 
age 0.75  pounds  per  square  foot.  These  weights  do  not 
include  the  interplane  wires  or  struts,  nor  the  fittings. 
The  total  weight  of  the  interplane  struts  of  the  JN-4B, 
the  Aviatic,  and  machines  of  similar  size  will  average 
from  28  to  30  pounds.  The  ailerons  will  weigh  about  12 
pounds  each. 

Weight  of  Motors.  There  is  a  considerable  difference 
in  the  weight  of  air-cooled  and  water-cooled  motors.  The 
water,  water  piping,  radiators  and  jackets  of  the  water- 
cooled  motors  adds  considerably  to  the  weight  of  the 
complete  power  plant.  The  mountings  are  heavier  for 
the  water-cooled  motors,  and  because  of  the  tandem 
arrangement  of  the  cylinders,  the  crankshaft  and  crank- 
case  weigh  more.  In  taking  the  bare  weight  of  the 
power  plant  all  of  the  accessories  must  be  included. 

In  the  following  table,  the  "bare  engine"  includes  the 
carbureter,  magneto,  and  necessary  integral  accessories, 
but  does  not  include  the  jacket  water,  mounting,  radiator, 
oil  in  base,  water  piping,  nor  controls.  Water-cooled 
motors  are  marked  by  (W)  and  air-cooled  by  (A). 
Rotary  air-cooled  are   (RA),  and  gallons   (G). 


306  WEIGHT  ESTIMATES 

WEIGHTS  OF  AERONAUTICAL  MOTORS. 

The  bare  radiator  will  weigh  from  0.48  to  0.56  pounds 
per  horsepower,  the  average  being  safe  at  0.52.  The 
water  contained  in  the  radiator  will  average  0.35  pounds 
per  horsepower.  The  weights  of  the  piping  and  the  water 
contained  therein  will  be  computed  separately.  The  cir- 
cular sheet  metal  cowl  used  over  the  rotary  cylinder  air- 
cooled  motor  is  equal  to  twice  the  square  root  of  the 
motor  weight,  according  to  Barnwell. 

Propeller  weight  varies  considerably  with  the  di- 
ameter, pitch,  etc.,  but  a  safe  rule  will  give  the  weight 
as  2.8  VH  where  H  =  horsepower.  The  tanks  will  weigh 
from  0.75  to  1.2  pounds  per  gallon  of  contents,  or  ap- 
proximately 1/5  the  weight  of  the  contents  when  com- 
pletely filled. 

Chassis  and  Wheel  Weight.  The  chassis  of  a  two- 
wheel  trainer  will  weigh  about  90  pounds  complete, 
although  there  are  chassis  of  training  machines  that 
weigh  as  much  as  140  pounds.  The  chassis  of  speed 
scouts  will  be  from  22  to  40  pounds  complete.  Tail  skids 
can  be  taken  at  from  6  to  8  pounds. 

Tangent  wire  wheels  complete  with  tires  are  about  as 
follows:  26x4^21  pounds;  26x5  =  28  pounds;  26x3 
=  14  pounds.  Ackerman  spring  spoke  wheels  are  es- 
timated as  follows:  20x4=17.5  pounds;  26x3  =  22 
pounds;  26x4  =  32  pounds;  30x4  =  35  pounds;  34x4 
=  45  pounds. 

Military  Loads.  A  20-mile  wireless  outfit  devised  by 
Capt.  Culver  weighed  40  pounds  with  storage  batteries, 
while  the  120-mile  outfit  weighed  60  pounds  with  a  180- 
watt  generator.  The  140-mile  U.  S.  A.  mule-back  wire- 
less of  1912  weighs  45  pounds.  The  "Blimp"  specifica- 
tions allow  250  pounds. 

The  Lewis  gun  as  mounted  on  the  "11"  Nieuport  weighs 
110    pounds,    including    mount,    gun    and    ammunition. 


WEIGHT  ESTIMATES 


307 


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308  WEIGHT  ESTIMATES 

Lewis  gun  bare  is  26  pounds.  The  Davis  6-pounder, 
Mark  IV,  weighs  103  pounds  with  mounting  but  with- 
out ammunition,  while  the  same  make  of  3-inch  12- 
pounder  weighs  238  pounds  under  the  same  conditions. 

Controls  and  Instruments.  The  Deperdussin  type  con- 
trols used  on  the  Curtiss  JN-4B  weigh  16  pounds  per 
control,  while  those  installed  in  the  Standard  H-3  weigh 
about  13  pounds.  An  average  of  15  pounds  per  control  is 
safe.  An  instrument  board  for  the  aviators'  cock-pit, 
fully  equipped,  weighs  from  20  to  24  pounds.  The  front, 
or  students'  instrument  board  will  average  10  pounds. 
Pyrene  extinguisher  and  brackets  =  7  pounds ;  Speak- 
ing tube  =  3  pounds ;  Oil  pressure  line  and  gage  =  3 
pounds;  Side  pockets  =  3  pounds;  Tool  kit —  10  pounds. 

Contr.ol  Surfaces.  The  rudder,  stabilizer,  fin,  and  ele- 
vator can  be  made  so  that  the  weight  will  not  exceed 
0.60  to  0.65  pounds  per  square  foot. 

General  Notes  on  Weight.  Before  starting  on  the 
weight  estimates  of  the  machine  the  reader  should  care- 
fully examine  the  tables  in  Chapter  II  which  give  the 
weights,  and  general  characteristics  of  a  number  of  mod- 
ern machines. 

Weights  and  Wing  Area.  When  the  weight  of  the 
machine  is  once  determined,  the  next  step  will  be  to 
determine  the  wing  area.  For  speed  scouts  or  very  large 
heavy  duty  machines  the  choice  of  a  wing  section  must 
be  very  carefully  considered.  For  the  speed  scout  sev- 
eral wings  giving  a  minimum  high  speed  resistance 
should  be  examined,  such  as  the  Eiffel  37  or  the  U.  S.  A.-l 
or  U.  S.  A.-6.  For  the  low-speed  aeroplane  to  be  de- 
signed for  great  lift,  a  number  of  sections  such  as  the 
U.  S.  A.-4  or  the  R.  A.  F.-3  should  be  tried  for  a  number 
of  speeds  and  angles.  For  training  machines  a  wing  of 
the  "x\ll  around"  type  such  as  the  R.  A.  F.-6  should  be 
adopted,  the  structural  characteristics  in  the  case  of  a 
trainer  having  an  important  bearing  on  the  subject.     If 


WEIGHT  ESTIMATES  309 

W  =  weight  of  the  machine  in  pounds,  V  =  low  speed 
in  miles  per  hour,  A  =  total  area  in  square  feet,  and 
Ky — lift  coefficient,  then  the  area  becomes  A=W/KyV-. 
Compensation  must  be  made  for  biplane  interference  for 
aspect  ratio,  and  stagger  as  previously  explained.  For 
an  ordinary  training  machine  with  the  usual  gap/chord 
ratio,  and  aspect  ratio,  the  correction  factor  of  0.85  may 
be  safely  employed. 

Example.  We  will  take  the  case  of  an  aeroplane  carry- 
ing a  personnel  load  of  340  pounds,  oil  and  gasoline  370 
pounds,  and  baggage  amounting  to  190  pounds,  instru- 
ments 100  pounds.  Total  live  load  will  be  1000  pounds. 
Taking  the  live  load  percentage  as  0.30,  the  total  load 
will  be  1000/0.30  =  3333  pounds.  If  the  low  speed  is  50 
miles  per  hour,  and  the  maximum  Ky  of  the  chosen  wing 
is  0.003  at  this  speed,  the  area  will  be  A  =  W/KyV^  = 
3333/0.003  X  (50  x  50)  =  444  square  feet.  Since  this  is  a 
biplane  with  a  correction  factor  of  0.85,  the  corrected  area 
will  be  :  444/0.85  =  523  square  feet.  The  unit  loading,  or 
weight  per  square  foot  will  be :  3333/523  =  6.36  pounds. 
The  corrected  area  includes  the  ailerons  and  the  part  of 
the  lower  wing  occupied  by  the  body. 

Empirical  Formula  for  Loading.  After  investigating 
a  large  number  of  practical  biplanes,  the  author  has  de- 
veloped an  expression  for  determining,  the  approximate 
unit  loading.  When  this  is  found,  the  approximate  area 
can  be  found  by  dividing  the  total  weight  by  the  unit 
loading.    This  gives  an  idea  as  to  the  area  used  in  practice. 

It  was  found  that  the  unit  loading  increased  with  the 
velocity  at  nearly  a  uniform  rate.  This  gave  an  average 
straight  line  formula  that  agreed  Very  closely  with  128 
examples.  If  V  =  Maximum  velocity  in  miles  per  hour, 
and  w  =  weight  per  square  foot,  then  the  unit  loading 
becomes : 

w  =  0.065V  —  0.25  for  the  average  case.  For  high 
speed  scouts  this  gives  a  result  that  is  a  trifle  low,  the 


310  WEIGHT  ESTIMATES 

formula  for  a  fast  machine  being  more  nearly  w  =  0.65V 

—  0.15,  for  speeds  over  100  miles  per  hour. 

A  two-seat  machine  of  average  size  weighs  2500 
pounds,  and  has  a  maximum  speed  of  90  miles  per  hour. 
Find  the  approximate  unit  loading  and  area.  The  load- 
ing becomes :  w  =  0.065  V  —  0.25  =  (0.065  x  90)  —  0.25 
=  5.6  pounds  per  square  foot.  The  approximate  area  will 
be:  2500/5.60  =  446  square  feet. 

If  the  above  machine  had  a  speed  of  110  miles  per  hour, 
the  formula  would  be  changed  for  the  high-speed  type 
machine,  and  the  loading  would  become : 

w  =  0.065V  —  0.15  =  (0.065  x  110)  —  0.15  =  7.00 
pounds  per  square  foot.  The  required  area  will  be : 
2500/7.0  =  372  square  feet.  When  the  unit  load  is  ap- 
proximately determined  in  this  way  it  is  a  very  simple 
matter  to  choose  the  wing  section  from  Ky  =  w/V'. 

Area  From  Live  Load  and  Speed.  By  a  combination 
of  empirical  formula  we  can  approximate  the  area  di- 
rectly. For  the  average  size  machine,  w^  0.065V  — 
0.25.  And  the  total  weight  W  =  U/0.32  where  U  is  the 
useful  or  live  load.    Since  A  =  W/w,  then  A  =  U/0.65 V 

—  0.25)  x  0.32  =  U/0.021  V  —  0.08. 

Thus  if  an  aeroplane  travels  at  90  miles  per  hour  and 
has  carried  a  useful  load  of  800  pounds  (including  gas  and 
oil),  the  approximate  area  is:  A^U/0.021V  =  0.08  = 
800/  (0.021  X  90)  —  0.08  =  442  square  feet.  This  as- 
sumes that  the  useful  load  is  0.32  of  the  total  load  and 
that  the  speed  is  less  than  100  miles  per  hour. 


CHAPTER  XV. 
BALANCE  AND  STABILITY. 

Elements  of  Stability.  When  we  balance  a  board  on  a 
fulcrum  so  that  it  stands  in  a  perfectly  horizontal  posi- 
tion, the  board  is  said  to  be  "In  equilibrium,"  or  is  sup- 
ported at  its  "Center  of  gravity."  There  is  only  one 
point  at  which  a  body  will  balance,  and  this  point  is  at 
the  center  of  gravity  or  "C.  G."  In  an  aeroplane,  the 
combined  mass  of  the  body,  motor,  wings,  fuel,  chassis, 
tail  and  live  load  has  a  center  of  gravity  or  a  balancing 
point  at  which  the  lift  must  be  applied  if  the  machine 
is  to  rest  in  equilibrium.  When  the  center  of  Hft  (or 
center  of  pressure)  does  not  pass  through  the  center  of 
gravity  of  the  aeroplane,  some  other  force  must  be  ap- 
plied to  overcome  the  unbalanced  condition.  When  the 
machine  is  unbalanced  in  a  fore  and  aft  direction  with 
the  tail  low,  a  force  must  be  applied  by  the  elevator  flaps 
that  is  opposite  and  equal  to  the  moment  of  the  un- 
balanced forces.  An  aeroplane  is  stable  when  it  is  bal- 
anced in  such  a  way  that  it  returns  to  a  state  of 
equilibrium  after  meeting  with  a  disturbance. 

When  disturbed,  a  stable  body  does  not  usually  return 
instantly  to  its  position  of  equilibrium,  but  reaches  it 
after  a  series  of  decreasing  oscillations.  The  heavier  the 
body,  and  the  more  compact  its  form,  the  longer  will  it 
oscillate  about  its  fulcrum  before  coming  to  rest.  By 
arranging  broad  surfaces  at  the  ends  of  the  oscillating 
body,  a  portion  of  the  energy  will  be  expended  in  cre- 
ating air  currents,  and  the  motion  will  be  readily 
"damped  out."    If  the  damping  effect  is  so  great  that  the 

311 


312  STABILITY 

body  does  not  swing  back  after  once  reaching  the  posi- 
tion of  equilibrium,  the  body  is  said  to  be  "dead  beat," 
or  "dynamically  stable."  There  is  a  great  difference 
between  the  static  forces  that  tend  to  return  the  body 
to  a  position  of  equilibrium  and  the  dynamic  retarding 
forces  that  tend  to  damp  out  the  oscillations.  Usually,  a 
body  with  excessive  static  stability  is  far  from  being 
stable  in  a  true  sense,  since  such  a  body  tends  to  oscillate 
longer,  and  more  violently,  than  one  in  which  the  static 
restoring  forces  are  not  so  strongly  marked.  A  body 
may  be  statically  but  not  dynamically  stable,  but  a 
dynamically  stable  body  must  of  necessity  be  statically 
stable. 

Static  stability  in  calm  air  is  determined  by  the  loca- 
tion of  the  center  of  gravity,  the  center  of  lift,  the  center 
of  propeller  thrust,  the  center  of  area  of  the  surfaces, 
and  the  center  of  the  forward  resistance.  The  forces  act- 
ing through  these  centers  are:  (1)  The  weight;  (2)  The 
lifting  force;  (3)  The  propeller  thrust;  (4)  The  resist- 
ance. The  weight  and  lift  are  vertical  forces  equal  and 
opposite  in  direction.  The  thrust  and  resistance  are  hori- 
zontal forces,  also  equal  and  opposite  in  direction.  When 
all  of  these  forces  intersect  at  a  common  point,  they  will 
completely  neutralize  one  another  and  the  body  will  be 
in  equilibrium. 

Dynamic  stability  is  attained  by  the  use  of  large  damp- 
ing surfaces  such  as  the  stabilizer  surface,  fins,  and  the 
elevator.  These  act  to  kill  the  oscillations. set  up  by  the 
static  righting  couples  or  forces.  Without  suitable  damp- 
ing surfaces  the  machine  would  soon  be  out  of  control  in 
gusty  weather  since  successive  wind  gusts  will  act  to  in- 
crease the  oscillations  of  the  righting  forces  until  the 
machine  will  turn  completely  over.  On  the  other  hand, 
an  aeroplane  can  be  too  stable  and  therefore  difficult  to 
steer  or  control  in  gusts  because  of  its  tendency  toward 
changing  its  attitude  with  every  gust  in  order  to  restore 


STABILITY  313 

its  equilibrium.  A  machine  should  only  be  partially 
stable,  and  the  majority  of  pilots  are  firmly  set  against 
any  form  of  mechanical  or  inherent  control.  No  matter 
how  simple  the  method,  mechanical  control  always  in- 
troduces a  certain  amount  of  mechanism  that  may  go 
wrong.  The  question  of  stabihty  has  already  been  solved 
to  a  sufficient  extent. 

A  disturbance  that  simply  changes  the  direction  of 
travel  is  not  considered  an  unstable  force  since  it  normal- 
ly does  not  tend  to  endanger  the  machine.  Nearly  any 
machine,  equipped  with  any  possible  form  of  control 
apparatus,  tends  to  change  its  direction  when  being 
righted. 

Axes  of  Stability.  An  aeroplane  has  six  degrees  of 
freedom  or  motion.  Three  are  of  translation  or  straight 
line  motion,  and  three  are  of  rotation  about  rectangular 
axes.  It  can  travel  forward  in  a  straight  line,  rise  and 
fall  in  a  vertical  plane,  or  skid  sidewise.  When  it  rolls 
from  side  to  side  about  the  fore  and  aft  axis  (X  axis) 
it  is  laterally  unstable.  When  pitching  up  and  down  in 
a  fore  and  aft  direction,  and  around  an  axis  parallel  with 
the  length  of  the  wings  (Y  axis),  the  machine  is  said  to 
be  longitudinally  unstable.  When  swinging  or  "Yaw- 
ing" from  right  to  left  about  a  vertical  axis  (Z  axis)  it  is 
unstable  in  "Yaw." 

Rolling  is  resisted  by  the  ailerons,  pitching  by  the  ele- 
vators and  stabilizer,  and  yawing  by  the  vertical  direc- 
tional rudder.  Lateral  oscillation  are  damped  out  by  the 
wing  surfaces  and  by  vertical  surfaces  or  "Fins."  Longi- 
tudinal oscillations  are  damped  mostly  by  the  stabilizer 
and  elevator  surfaces.  Directional  or  yawing  vibrations 
are  corrected  by  the  damping  action  of  the  vertical  tail 
fin,  vertical  rudder  and  the  sides  of  the  body,  the  latter 
also  serving  to  damp  out  longitudinal  vibrations.  On  an 
absolutely  calm  day,  the  pilot  can  shut  oft  the  motor 
and   glide   down   without   touching  the   controls   if   the 


314  STABILITY 

machine  is  longitudinally  stable.  The  glide  generally 
starts  with  a  few  pitching  oscillations,  but  these  grad- 
ually are  damped  out  by  the  tail  as  soon  as  the  machine 
picks  up  its  natural  gliding  angle  and  speed,  and  from 
this  point  it  will  continue  without  oscillating. 

The  Spiral  and  Nose  Dive.  There  are  two  forms  of 
instability  that  have  not  yet  been  fully  corrected,  and 
both  are  highly  dangerous.  One  of  these  is  known  as 
the  "spiral  dive"  or  nose  spin,  and  the  other  as  the 
straight  nose  dive.  The  aeroplane  in  a  spiral  nose  dive 
rotates  rapidly  about  a  vertical  axis  during  the  dive. 
Spiral  instability  resulting  from  lateral  instability,  can 
be  minimized  by  decreasing  the  area  of  the  vertical  rud- 
der and  by  the  proper  placing  of  fins  so  that  there  is  not 
so  great  an  excess  of  vertical  area  to  the  rear  of  the  C.  G. 

The  covered-in  body  acts  as  a  fin  and  will  be  pro- 
ductive of  spiral  instability  if  the  area  is  not  properly 
distributed.  In  the  majority  of  cases  the  rear  of  the  body 
is  equivalent  to  a  large  fin  placed  to  the  rear  of  the  C.  G. 
A  fin  above  the  G.  G.  tends  to  reduce  all  spiralling. 

Stability  and  Speed.  An  aeroplane  in  straight  hori- 
zontal flight  must  be  driven  at  such  an  angle,  and  such  a 
speed,  that  the  weight  is  just  sustained.  To  be  inherently 
stable  the  machine  must  always  tend  to  increase  its  speed 
by  diving  should  the  power  be  cut  off  in  any  way.  An 
aeroplane  that  does  not  tend  to  increase  its  speed  in  this 
way,  "Stalls"  or  becomes  out  of  control.  Any  machine 
that  will  automatically  pick  up  its  gliding  angle  after  the 
propeller  thrust  has  ceased  is  at  least  partially  inherently 
stable,  and  if  it  does  not  possess  this  degree  of  stability, 
other  forms  of  stability  are  practically  worthless.  The 
machine  having  the  smallest,  flattest  gliding  angle  is  nat- 
urally safest  in  cases  of  power  failure,  and  hence  the  glid- 
ing angle  is  somewhat  related  to  the  subject  of  stability. 

The  longitudinal  stability  decreases  with  a  decrease 
in  the  speed,  the  fore  and  aft  vibrations  becoming  more 


STABILITY 


315 


rapid  due  to  the  decreased  effect  of  the  tail  surfaces,  and 
to  the  reduction  of  wing  Hft.  InstabiHty  at  low  speeds  is 
common  to  all  aeroplanes,  whether  inherently  stable  or 
not,  and  at  a  certain  critical  speed  the  machine  becomes 
absolutely  unstable  in  a  dynamic  sense.  If  a  machine 
is  to  be  stable  at  low  speeds,  it  must  not  fly  at  too  great 


A  Spanish  Aeroplane  Using  a  Peculiar  Form  of  Upper  Fin.     These  Fins  Also 
Perform  the  Duty  of  Vertical  Rudders  as  Well  as  Acting  as  Stabilizers. 

an  angle  of  incidence  at  these  speeds,  and  it  should  have 
a  very  large  tail  surface  acting  at  a  considerable  distance 
from  the  wings.  Hunsaker  states  that  the  lowest  speed 
should  not  require  more  than  80  per  cent  of  the  total  lift 
possible. 

Inertia  or  Flywheel  EfiFect.  The  principal  weights 
should  be  concentrated  as  nearly  as  possible  at  the  center 
of  gravity.     Weights  placed  at  extreme  outer  positions. 


316  STABILITY 

as  at  the  wing  tips,  or  far  ahead  of  the  wings,  tend  to 
maintain  oscillations  by  virtue  of  their  flywheel  effect. 
)  The  measure  of  this  inertia  or  flywheelage  is  known  as 
!  the  "Moment  of  Inertia"  and  is  the  sum  of  the  products 
of  all  the  masses  by  the  squares  of  their  distances  from 
the  center  of  gravity.  A  great  amount  of  inertia  must  be 
met  by  a  large  damping  surface  or  control  area  if  the 
vibrations  are  to  be  damped  out  in  a  given  time.  In  twin- 
motored  aeroplanes  the  motors  should  be  kept  as  close 
to  the  body  as  the  propellers  will  permit. 

Wind  Gusts  and  Speed.  A  machine  flying  at  high 
speed  is  less  affected  by  wind  gusts  or  variations  in  den- 
sity than  a  slow  machine,  since  the  disturbing  currents 
are  a  smaller  percentage  of  the  total  speed.  In  addition, 
a  high  speed  results  in  smaller  stresses  due  to  the  gusts. 

Gyroscopic  Instability.  The  motor  gyroscopic  forces 
do  not  affect  the  stability  of  a  machine  to  any  great  ex- 
tent, and  in  twin  motored  aeroplanes  the  gyroscopic  ac- 
tion of  the  propellers  is  almost  entirely  neutralized.  At 
one  time  the  gyroscopic  torque  was  blamed  for  every 
form  of  instability,  but  on  investigation  it  was  found 
that  the  practical  effect  was  negligible. 

Instability  Due  Power  Plant.  The  power  plant  affects 
stability  in  a  number  of  ways.  The  thrust  of  the  pro- 
peller may  cause  a  fore  and  aft  moment  if  the  center  line 
of  thrust  does  not  pass  through  the  center  of  resistance. 
This  causes  the  machine  to  be  held  head  up,  or  head 
down,  according  to  whether  the  line  of  thrust  is  below 
or  above  the  C.  G.  If  the  propeller  thrust  tends  to  hold 
the  head  up  in  normal  flight,  the  machine  will  tend  to 
dive,  and  assume  its  normal  gliding  velocity  with  the 
power  off,  hence  this  is  a  condition  of  stability.  With 
the  effect  of  the  thrust  neutral,  or  with  the  thrust  passing 
through  the  center  of  resistance,  the  machine  will  not 
tend  to  maintain  the  speed,  and  hence  it  is  likely  to  stall 
unless  immediately  corrected  by  the  pilot.    With  the  line 


STABILITY 


317 


of  thrust  above  the  C.  G.,  the  stall  effect  is  still  further 
increased  since  with  this  arrangement  there  is  a  very 
decided  tendency  for  the  machine  to  nose  up  and  increase 
the  angle  of  incidence  when  the  power  is  cut  oft*. 

The  slip  stream  of  the  propeller  has  a  very  decided 
effect  on  the  tail  surfaces,  these  being  much  more  effec-    I 
tive  when  the  propeller  slip  stream  passes  over  them. 


Steel  Elevator  and  Rudder  Construction  Used  on  a  European  Machine.  The 
Elevators  Also  Act  as  Stabilizers,  the  Entire  Surface  Turning  About 
the   Tube    Spar. 

With  lifting  tails,  or  tails  that  normally  carry  a  part  of 
the  load,  the  stoppage  of  the  slip  stream  decreases  the 
lift  of  the  tail  and  consequently  tends  to  stall  the 
machine.  Non-lifting  tails  should  be  arranged  so  that 
the  slip  stream  strikes  down  on  the  upper  surface.  This' 
tends  to  force  the  tail  down,  and  the  head  up  in  normal; 
flight,  and  when  the  power  ceases  the  tail  will  be  re-l 
lieved  and  there  will  be  an  automatic  tendency  toward] 
diving  and  increase  in  speed.     On  a  twin  aeroplane,  a 


318  STABILITY 

similar  effect  is  obtained  by  making  the  upper  tips  of 
both  propellers  turn  inwardly.  The  air  is  thus  thrown 
down  on  the  tail. 

With  a  single  motor,  the  "Torque  tends  to  turn  the 
aeroplane  in  a  direction  opposite  to  the  rotation  of  the 
propeller.  Lateral  stability  is  thus  interfered  with  when 
the  motor  is  cut  off  or  reduced  in  speed.  With  right- 
hand  propeller  rotation,  for  example,  the  machine  will  be 
turned  toward  the  left,  forcing  the  left  tip  down.  To 
maintain  a  horizontal  attitude,  the  left  aileron  must  be 
held  down  by  an  amount  just  sufficient  to  overcome  the 
torque.  In  some  machines  one  wing  tip  is  given  a  perma- 
nent increase  in  incidence  so  that  the  down  seeking  tip 
is  given  permanent  additional  lift. 

Lateral  Stability.  WTien  an  aeroplane  is  turned  sharp- 
ly in  a  horizontal  plane,  or  *'Yaws,"  the  outer  and  faster 
moving  wing  tip  receives  the  greater  lift,  and  a  lateral 
rolling  moment  is  produced  about  the  fore  and  aft  axis. 
In  the  opposite  condition,  a  lateral  rolling  moment  tends 
to  yaw  or  to  throw  the  aeroplane  off  a  straight  course. 
Below  a  certain  critical  speed,  the  lateral  or  rolling  oscil- 
lations increase  in  amplitude,  with  a  strong  tendency 
to  side  slip,  skid  or  spiral.  The  tail  fin  or  rudder  retards 
the  tail  velocity  in  a  side  slip,  and  thus  turns  the  slipping 
or  skidding  machine  into  a  vertical  spiral  or  spinning 
nose  dive.  This  spin  increases  the  angle  of  bank  and 
hence  the  side  slip.  This  in  turn  increases  the  turning 
or  yawing  velocity,  and  the  spiral  starts.  This  tendency 
toward  a  spiral  dive  can  be  corrected  by  a  vertical  fin 
placed  forward,  and  above  the  center  of  gravity,  or  by 
raising  the  wing  tips.  An  upper  fin  of  this  type  will  give 
a  force  that  tends  to  break  up  the  bank  when  side  slip 
starts  and  thus  will  prevent  spinning. 

At  normal  speeds  the  rolling  is  damped  down  by  the 
wing  surfaces,  and  can  be  further  controlled  by  the  ap- 
plication  of  the   ailerons.     At  the   lower   critical   speed 


STABILITY 


319 


?^    OJ    aj 


x-2  2 


Bw5 
"on  «j  a 

^:^ 

■ail 

o  S  --> 


320  STABILITY 

when  the  machine  is  stalled,  one  wing  tip  has  no  more 
lift  than  the  other,  and  hence  the  damping  effect  of  the 
wings  and  the  action  of  the  ailerons  becomes  negligible. 

Dutch  Roll.  In  "Dutch  Roll,"  the  rolling  is  ac- 
companied by  an  alternate  yawing  from  right  to  left. 
This  is  aggravated  by  a  fin  placed  high  above  the  C.  G., 
and  hence  corrections  for  spiral  dive  conflict  with  correc- 
tions for  Dutch  roll.  The  rolling  is  accompanied  by  some 
side  slip,  and  the  motion  is  stable  providing  that  there  is 
sufficient  fin  in  the  rear  and  not  an  excessive  amount 
above  the  C.  G. 

Degree  of  Stability.  Excessive  stability  is  dangerous 
unless  the  control  surfaces  are  powerful  enough  to  over- 
come the  stable  tendency.  Since  a  stable  machine  always 
seeks  to  face  the  relative  wind,  it  becomes  difficult  to 
handle  in  gusty  weather,  as  it  is  continually  changing  its 
course  to  meet  periodic  disturbances.  This  is  aggravated 
by  a  high  degree  of  static  stability,  and  may  be  positively 
dangerous  when  landing  in  windy  weather. 

Control  Surfaces.  A  non-lifting  tail  must  give  no  lift 
when  at  a  zero  angle  of  incidence.  It  must  be  sym- 
metrical in  section  so  that  equal  values  of  lift  are  given 
by  equal  positive  and  negative  angles  of  incidence. 
Square  edged,  flat  surfaces  are  not  desirable  because  of 
their  great  resistance.  A  double  cambered  surface  is 
suitable  for  such  controls  as  the  stabilizer,  elevator  and 
rudder.  It  has  a  low  resistance,  permits  of  strong  in- 
ternal spars,  and  is  symmetrical  about  the  line  of  the 
chord.  Some  tails  are  provided  with  a  cambered  top  and 
a  flat  bottom  surface  so  that  the  down  wash  of  the  wings 
is  neutralized.  Under  ordinary  conditions  this  would  be 
an  unsymmetrical  lifting  surface,  but  when  properly 
adapted  to  the  wings  the  lifting  effect  is  completely 
neutralized  by  the  down  v^ash. 

The  curvature  of  the  section  should  be  such  that  the 
movement  of  the  center  of  pressure  is  as  small  as  possi- 


STABILITY  321 

ble.  With  a  small  movement  of  the  center  of  pressure, 
the  surface  can  be  accurately  balanced  and  hinged  on 
the  center  of  pressure  line.  It  is  desirable  to  have  the 
maximum  thickness  of  section  at,  or  near  to  the  C.  P., 
so  that  a  deep  spar  can  be  used  for  the  support  of  the 
hinge  system.  Usually  the  movement  of  the  control  sur- 
faces is  limited  to  an  angle  of  30  degrees  on  either  side 
of  the  center  line,  as  the  lift  of  all  surfaces  start  to  de- 
crease after  this  point  is  reached.  The  surface  move- 
ment should  be  limited  by  the  maximum  lift  angle  of 
the  section  in  any  case,  since  an  accident  will  be  bound 
to  occur  if  they  are  allowed  movement  beyond  the  angle 
of  maximum  lift. 

In  locating  the  control  surfaces,  careful  attention 
should  be  paid  to  the  surrounding  air  conditions  so  that 
they  will  not  be  unduly  affected  by  the  wash-down  of  the 
wings  or  body.  The  effectiveness  of  the  tail  surfaces  is 
very  much  reduced  by  bringing  them  close  to  the  wings, 
and  the  lift  is  always  reduced  by  the  wash  of  a  covered 
fuselage. 

The  wash-down  effect  of  the  wings  on  the  tail  is  pro- 
portional to  the  chord  and  not  to  the  span,  and  for  this 
reason  an  increase  in  span  does  not  always  necessitate 
an  increase  in  the  length  of  the  body.  An  adequate  damp- 
ing effect  requires  a  large  surface  at  the  end  of  a  long 
lever  arm. 

Balancing  the  Aeroplane.  Figs.  1  to  6  show  the  prin- 
ciples involved  in  the  balancing  of  the  aeroplane.  In 
Fig.  1  a  number  of  weights  r-2'-3'  and  5M  are  supported 
on  a  beam,  the  load  being  balanced  on  the  fulcrum  point 
M.  The  load  2'  being  directly  over  the  fulcrum,  has  no 
influence  on  the  balance,  but  load  V  at  the  left  tends  to 
turn  the  mass  in  a  left-hand  direction,  while  3'  and  5M 
tend  to  give  it  a  right-hand  rotation.  This  turning 
tendency  depends  upon  the  weights  of  the  bodies  and  their 
distance   from   the   fulcrum.     The   turning  tendency   or 


322 


STABILITY 


"Moment"  is  measured  by  the  product  of  the  weight  and 
the  distance  from  the  fulcrum.  If  weight  V  should  be  10 
pounds,  and  its  distance  A'  from  the  fulcrum  should  be 
20  inches,  then  it  would  cause  a  left-hand  moment  of 
10x20^200  inch  pounds.  If  the  system  is  to  be  in 
balance,  then  the  left-hand  moment  of  V  should  be  equal 
to  the  sum  of  the  moments  of  3'  and  5M.  Thus:  1'  x  A 
=  (3'xB)  +  (SMxC). 

The  application  of  this  principle  as  applied  to  a  mono- 
plane is  shown  by  Fig.  4,  in  which  X-X  is  the  center  of 


Figs.   1-6.     Methods  of  Balancing  an   Aeroplane  About  Center  of  Lift. 

pressure  or  lift.  The  center  of  lift  corresponds  to  the 
fulcrum  in  Fig.  1,  and  the  weights  of  the  aeroplane 
masses  and  their  distance  from  the  center  of  lift  are 
shown  by  the  same  letter  as  in  Fig.  1.  The  engine  V  is 
at  the  right  of  the  C.  P.  by  the  distance  A,  while  the  fuel 
tank  2  is  placed  on  the  C.  P.  in  the  same  way  that  the 
weight  2'  in  Fig.  1  is  placed  directly  over  the  fulcrum. 
By  placing  the  tank  in  this  position,  the  balance  is  not 
affected  by  the  emptying  of  the  fuel  since  it  exerts  no 
moment.  The  chassis  G  acting  through  the  distance  E 
is  in  the  same  direction  as  the  engine  load.  The  body  5 
with  its  center  of  gravity  at  M  acts  through  the  distance 
C,  while  the  weight  of  the  pilot  3  exerts  a  right-hand 
moment  with  the  lever  arm  length  B.     If  the  moments 


STABILITY  323 

of  all  these  weights  are  not  in  equilibrium,  an  additional 
force  must  be  exerted  by  the  tail  V. 

Fig.  2  shows  an  additional  weight  4'  that  corresponds 
to  the  weight  of  the  passenger  4  in  Fig.  5.  This  tends 
to  increase  the  right  turning  moment  unless  the  fulcrum 
is  moved  toward  the  new  load.  In  Fig.  2  the  fulcrum  M 
remains  at  the  same  point  as  in  Fig.  1,  hence  the  system 
requires  a  new  force  P'  acting  up  at  the  end  of  the  beam. 
If  ihe  load  was  in  equilibrium  before  the  addition  of  4', 
then  the  force  P'  must  be  such  that  P'xT'  =  4'xD'. 
In  the  equivalent  Fig.  5,  the  center  of  gravity  has  moved 
from  its  former  position  at  S  to  the  new  position  at  R, 
the  extent  of  the  motion  being  indicated  by  U.  To  hold 
this  in  equilibrium,  an  upward  force  P  must  be  exerted 
by  the  elevator  at  Y,  the  lever  arm  being  equal  to 
(T  +  U). 

Fig.  6  shows  the  single-seater,  but  under  a  new  condi- 
tion, the  center  of  pressure  having  moved  back  from 
X-X  to  Z.  To  hold  the  aeroplane  in  equilibrium,  a  down- 
ward force  must  be  provided  by  the  tail  V  which  will 
cause  a  right-hand  moment  equal  to  the  product  of  the 
entire  weight  and  the  distance  U.  For  every  shift  in 
the  center  of  pressure,  there  must  be  a  corresponding 
moment  provided  by  the  elevator  surface.  The  condition 
is  shown  by  the  simple  loaded  beam  of  Fig.  3.  In  this 
case  the  fulcrum  has  been  moved  from  M  to  N,  a  distance 
equal  to  the  center  of  pressure  movement  in  Fig.  6.  This 
requires  a  downward  force  P'  to  maintain  equilibrium. 

Center  of  Pressure  Calculation.  Fig.  7  is  a  diagram 
showing  the  method  of  calculating  the  center  of  gravity. 
The  reference  line  R  is  shown  below  the  elevators  and  is 
drawn  parallel  to  the  center  of  pressure  line  W-W,  the 
latter  line  being  assumed  to  pass  through  the  center  of 
gravity.  The  line  R  may  be  located  at  any  convenient 
point,  as  at  the  propeller  flange  or  elsewhere,  but  for 
clearness  in  illustration  it  is  located  to  the  rear  of  the 


324 


STABILITY 


aeroplane.  The  weight  of  each  item  is  multiplied  by  the 
distance  of  its  center  of  gravity  from  the  line  R,  these 
products  are  added,  and  the  sum  is  then  divided  by  the 
total  weight  of  the  machine.  The  result  of  this  division 
gives  the  distance  of  the  center  of  gravity  from  the  line 
R.  Thus,  if  the  center  of  gravity  of  the  body  (11)  is 
located  at  (10),  then  the  product  of  the  body  weight  mul- 
tiplied by  the  distance  B  will  give  the  moment  of  the 
body  about  the  line  R.  The  weight  of  the  motor  (2)  mul- 
tiplied by  the  distance  F  gives  the  moment  of  the  motor 
about  R,  and  so  on  through  the  list  of  items. 


CENTER  OF  GRAVITY  TABLE. 

Mark                Name  of                Weight  of  Distance  from  Product 

No.                      Item                          Item  Line  R-R 

1  Propeller 10  lbs.  G  ^  230"  2,300 

2  Motor  and  Radiator..  .     400  lbs.  F=190"  76,000 
4-6  Wing     Surfaces 200  lbs.  E=180"  36,000 

5  Chassis     75  lbs.  190"  14,250 

8  Pilot  and  Seat 170  lbs.  C=150"  25,800 

9  Fuel    Tank 120  lbs.  D=170"  20,400 

1 1  Fuselage    300  lbs.  B  -=  140"  42,000 

12  Stabilizer     40  lbs.  A  =    50"  2,000 

15  Horizontal    Rudder...        10  lbs.  K=    20"  200 

16  Elevator     20  lbs.  K  =    20"  400 

Totals 1,375  219,350  Total  Product 


The  distance  of  the  center  of  gravity  (or  center  of 
pressure)  from  the  reference  line  R  is  given  by  H  +  K. 
This  gives  the  numerical  value  219350/1375=159.6 
inches.  Thus  if  we  measure  159.6  inches  from  R  toward 
the  wings  we  will  have  located  the  center  of  gravity.  The 
location  of  the  C.  G.  can  be  changed  by  shifting  the 
weights  of  the  motor,  passenger,  or  other  easily  moved 
items.  In  any  case,  the  C.  G.  should  lie  near  the  center  of 
pressure. 

Tail  Lever  Arms.  The  effective  damping  moment 
exerted  by  the  fixed  stabilizer  surface  (12)  will  be  the 
product  of  its  area  by  the  distance  (I),  measured  from 
the  center  of  pressure  of  the  wing  to  the  center  of 
pressure  of  the  stabilizer.    The  lever  arm  of  the  elevator 


STABILITY 


325 


is  the  distance  (H)  measured  from  the  centers  of  pressure 
as  before. 

Resultant  Forces  and  Moments  in  Flight.  The  aero- 
plane is  in  equilibrium  when  all  of  the  forces  pass  through 
a  common  center,  as  shown  by  Fig.  8.  In  this  figure  the 
lift  (L),  the  weight  (W),  the  line  of  propeller  thrust  (T), 
and  the  resistance    (R)   all  pass  through  the  center  of 


.^ 


u  u.  o 





Fig.  7.     Method  of  Determining  the  Center  of  Gravity  of  an  Aeroplane. 


gravity  shown  by  the  black  dot  C.  G.  There  are  no 
moments  and  hence  no  correction  is  needed  from  the  ele- 
vator (T).  In  Fig.  9,  the  thrust  and  resistance  pass 
through  the  center  of  gravity  as  before,  but  the  center  of 
lift  (L)  does  not  pass  through  the  center  of  gravity,  the 
distance  betwen  the  two  being  indicated  by  (n).  This 
causes  a  moment,  the  length  of  the  lever  arm  (n)  being 
effective  in  giving  a  right-hand  rotation  to  the  body.     If 


326  STABILITY 

horizontal  flight  is  to  be  had  this  must  be  resisted  by  the 
upward  elevator  force  (E). 

In  Fig.  10,  the  lift  passes  through  the  center  of  gravity, 
but  the  line  of  resistance  lies  below  it  by  the  amount  (m). 
The  thrust  (T)  tends  to  rotate  the  machine  in  a  left- 
handed  direction.  The  elevator  must  exert  a  downward 
force  (e)  to  resist  the  moment  caused  by  (m).  This  is  a 
bad  disposition  of  forces,  as  the  machine  would  tend  to 
stall  or  tail-dive  should  the  propeller  thrust  cease  for 
even  an  instant.  The  stability  of  Figs.  8  and  9  would  not 
be  affected  by  the  propeller  thrust,  as  it  passes  through 
the  C.  G.  in  both  cases.  In  Fig.  11,  the  center  line  of 
thrust  is  below  the  line  of  resistance  (R),  so  that  the 
thrust  tends  to  hold  the  nose  up.  Should  the  motor  fail 
in  this  case,  the  nose  would  drop  and  the  machine  would 
start  on  its  gliding  angle  and  pick  up  speed. 

In  Fig.  12  none  of  the  forces  intersect  at  a  common 
point,  the  lift  and  weight  forming  a  right-handed  couple, 
while  the  thrust  (T)  and  the  resistance  (R)  form  a  left- 
handed  couple  that  opposes  the  couple  set  up  by  the 
weight  and  lift  forces.  If  the  thrust-resistance  couple 
can  be  made  equal  to  the  lift-weight  couple,  the  aero- 
plane will  be  in  equilibrium  and  will  need  no  assistance 
from  the  elevator.  As  the  weights  in  the  aeroplane  are 
all  located  at  different  heights,  it  is  necessary  to  obtain 
the  center  of  gravity  of  all  the  loads  in  a  vertical  plane 
as  well  as  horizontally.  Thus  in  Fig.  13  the  line  C.  G.  is 
the  center  of  gravity  of  the  engine  weight  (1),  the  wing 
weight  (2),  the  pilot's  weight  (3),  the  chassis  weight 
(4),  the  fuselage  weight  (5),  and  the  fuel  tank  weight 
(6).  The  Hne  C.  G.  is  the  effective  center  of  all  these 
loads,  and  is  calculated  by  taking  the  products  of  the 
weights  by  the  distance  from  a  reference  line  such  as 
R-R.  The  center  of  resistance  is  the  effective  center  of 
all  the  resistance  producing  items  such  as  the  wings, 
body,  struts,  chassis,  etc. 


STABILITY 


327 


A  suggestion  of  the  method  employed  in  obtaining  the 
center  of  resistance  is  shown  by  Fig.  14,  the  center  line 
of  resistance  R-R  being  the  resultant  of  the  wing  resist- 
ance (D),  the  body  resistance  (B),  and  the  chassis  re- 
sistance (C).     It  will  be  noted  that  the  wing  resistance 


Figs.  8-15.     Forces   Affecting  the   Longitudinal    Stability    of   an   Aeroplane. 


of  biplane  wings  (W-W)  does  not  lay  midway  between 
the  wings  but  rather  closer  to  the  upper  wing,  as  shown 
by  (E).  This  is  due  to  the  upper  wing  performing  the 
greater  part  of  the  lift.  In  locating  the  center  of  resist- 
ance, the  resistance  forces  are  treated  exactly  like  the 
weights  in  the  C.  G.  determination.  Each  force  is  mul- 
tiplied by  its  distance  from  a  horizontal  reference  line, 


328  STABILITY 

and  the  sum  of  the  products  is  divided  by  the  total 
resistance.  As  shown,  the  center  of  resistance  R-R 
passes  through  the  center  of  gravity  C.  G.  The  center  of 
pressure  Hne  X-X  also  contains  the  center  of  resistance. 

A  staggered  biplane  cell  is  shown  by  Fig.  15,  the  cen- 
ter of  pressure  of  the  upper  and  lower  wings  being  con- 
nected by  the  line  X-X  as  before.  The  center  of  resist- 
ance of  the  pair  is  shown  at  (D),  where  it  is  closer  to  the 
upper  wing  than  to  the  lower.  A  vertical  line  Y-Y 
dropped  through  the  center  of  resistance  gives  the  loca- 
tion of  the  center  of  lift.  As  shown,  the  center  of  lift  is 
brought  forward  by  the  stagger  until  it  is  a  distance  (g) 
in  front  of  the  leading  edge  of  the  lower  wing.  The  cen- 
ter of  lift  and  the  center  of  resistance  both  lie  on  a  line 
connecting  the  center  of  pressure  of  the  upper  and  lower 
wings. 

Calculation  of  Control  Surfaces.  It  is  almost  impossi- 
ble to  give  a  hard  and  fast  rule  for  the  calculation  of  the 
control  surfaces.  The  area  of  the  ailerons  and  tail  sur- 
faces depends  upon  the  degree  of  stability  of  the  main 
wings,  upon  the  moment  of  inertia  of  the  complete 
machine,  and  upon  the  turning  moments.  If  the  wings 
are  swept  back  or  set  with  a  stagger-decalage  arrange- 
ment, they  will  require  less  tail  than  an  orthogonal  cell. 
All  of  these  quantities  have  to  be  worked  out  differently 
for  every  individual  case. 

Aileron  Caculations.  The  ailerons  may  be  used  only 
on  the  upper  wing  (2  ailerons),  or  they  may  be  used  on 
both  the  upper  and  lower  wings.  When  only  two  are 
used  on  the  upper  wing  it  is  usually  the  practice  to  have 
considerable  overhang.  When  the  wings  are  of  equal 
length  either  two  or  four  ailerons  may  be  used.  Roughly, 
the  ailerons  are  about  one-quarter  of  the  wing  span  in 
length.  With  a  long  span,  a  given  aileron  area  will  be 
more  effective  because  of  its  greater  lever  arm. 

If  a  =  area  of  ailerons,  and  A  =  total  wing  area  in 


STABILITY  329 

square  feet,  with  S  =  wing  span  in  feet,  the  aileron  area 
becomes:  a  =  3.2A/S.  It  should  be  borne  in  mind  that 
this  appHes  only  to  an  aeroplane  having  two  ailerons  on 
the  upper  wing,  since  a  four-aileron  type  usually  has 
about  50  per  cent  more  aileron  area  for  the  same  wing 
area  and  wing  span.  For,  example,  let  the  wing  span  be 
40  feet  and  the  area  of  the  wings  be  440  square  feet,  then 
the  aileron  area  will  be:  a  =  3.2A/S  =  3.2  x  440/40  = 
35.2  square  feet.  If  four  ailerons  were  employed,  two  on 
the  upper  and  two  on  the  lower  wing,  the  area  would  be 
increased  to  1.5  x  35.2  =  52.8  square  feet.  As  an  example 
in  the  sizes  of  ailerons,  the  following  table  will  be  of 
interest: 

AILERON  SIZES. 

Name  of                    Year        Wing      Wing  No.  of  Area  of  Total    Length 
Machine                     Tvpe         Span        Area    Ailers    1  Ai'er    Area     Ailer 
L.W.F.     (Amer.) 1916         46'6"         490         2         19.0         38  0         10'3" 

38'8" 
Eastern    (Amer.) 1916         42'0"         420         2         23.0         46.0         lO'O" 

39'0" 
Mavo    (Amer.) 1915         38'0"         375         4         16.0         64.0  S'O" 

38'0" 
London-Prov 1916         37'0"         350         4  13.5  54.0  S'O" 

37'0" 
Albatros    (Ger.) 1915         43'3"         458         2         16.2         32.4  9'3" 

37'3" 

Standard  (Amer.) 1916    40'1"    520    4    15.5    62.0     8'4" 

H-3  40'1" 
Nieuport  (Scout) 1916    24'6"    145    2     7.0    14.0     5'9" 

23'0" 
Curtiss  (Baby) 1916    20'0"    155    2     7.0    14.0     4'9" 

18'0" 

In  cases  where  the  upper  and  lower  spans  are  not 
equal,  take  the  average  span — that  is,  one-half  the  sum 
of  the  two  spans. 

Stabilizer  and  Elevator  Calculations.  These  surfaces 
should  properly  be  calculated  from  the  values  of  the 
upsetting  couples  and  moments  of  inertia,  but  a  rough 
rule  can  be  given  that  will  approximate  the  area.  If  a'  = 
combined  area  of  stabilizer  and  elevator  in  square  feet; 
L  =  distance  from  C.  P.  of  wings  to  the  C.  P.  of  tail  sur- 
face ;  A  =  Area  of  wings  in  square  feet,  and  C  =  chord 
of  wings  in  feet,  then, 

a'  =  0.51AC/L.    Assuming  our  area  as  430  square  feet, 


330  STABILITY 

the  chord  as  5.7  feet,  and  the  lever  arm  as  20  feet,  then, 
a'  =  0.51AC/L  =  0.51  x  430  x  5.7/20  =  62.5  square 
feet,  the  combined  area  of  the  elevators  and  stabilizer. 
The  relation  betwen  the  elevator  and  stabilizer  areas  is 
not  a  fixed  quantity,  but  machines  having  a  stabilizer 
about  20  per  cent  greater  than  the  elevator  give  good 
results.  In  the  example  just  given,  the  elevator  area  will 
be:  62.5/2.2:^28.41  square  feet,  where  2.2  is  the  constant 
obtained  from  the  ratio  of  sizes.  The  area  of  the 
stabilizer  is  obtained  from:  28.41x1.2  =  34.1  square 
feet. 

Negative  Stabilizers.  A  considerable  amount  of  in- 
herent longitudinal  stability  is  obtained  by  placing  the 
stabilizing  surface  at  a  slight  negative  angle  with  the 
wings.  This  angle  generally  varies  from  — 2°  to  — 6°. 
At  small  angles  of  wing  incidence  the  negative  angle  of 
the  tail  will  be  at  a  maximum,  and  acting  down  will 
oppose  further  diving  and  tend  to  head  the  machine  up. 
At  large  wing  angles,  the  tail  will  be  depressed  so  far 
that  the  tail  angle  will  become  positive  instead  of  nega- 
tive, and  thus  the  lift  on  the  tail  will  oppose  the  wings 
and  will  force  the  machine  to  a  smaller  angle  of  incidence. 
The  negative  angle  can  thus  be  adjusted  to  give  longi- 
tudinal stability  within  the  ordinary  range  of  flight 
angles. 

Stabilizer  Shapes  and  Aspect  Ratio.  Stabilizers  have 
been  built  in  a  great  number  of  different  shapes,  semi- 
circular, triangular,  elliptical,  and  of  rectangular  wing 
form.  Measured  at  the  rear  hinged  joint,  the  span  or 
width  of  the  stabiHzer  is  about  1/3  the  wing  span  for 
speed  scouts,  and  about  1/4  the  wing  span  for  the  larger 
machines.  Nearly  all  modern  machines  have  non-lifting 
tails,  or  tails  so  modified  that  they  are  nearly  non-lifting. 
Since  flat  plates  give  the  greatest  lift  with  a  small  aspect 
ratio,  and  hence  are  most  effective  when  running  over 
the  ground  at  low  speeds,  the  stabilizers  and  elevators 


STABILITY 


331 


are  of  comparatively  low  aspect.  In  general,  an  aspect 
ratio  of  3  is  a  good  value  for  the  stabilizer.  Vertical  rud- 
ders generally  have  an  aspect  ratio  of  1,  and  hence  are 
even  more  effective  per  unit  area  than  the  stabilizers. 
This  is  particularly  necessary  in  ground  running. 

Vertical    Rudders.     The    calculation    of    the    vertical 
rudders  must  take  the  moment  of  inertia  and  yawing 


Stick  Control  Used  on  the  Caudron  Biplane.     Wing  Warp  Is  Used  Instead  of 
Ailerons.      Back    and    Forth   Movement   Actuates    Elevator. 

moments  into  effect,  and  this  is  rather  a  complicated 
calculation  for  the  beginner.  As  an  approximation,  the 
area  of  the  rudder  can  be  taken  from  9  to  12  square  feet 
for  machines  of  about  40  feet  span,  and  from  5  to  8  square 
feet  for  speed  scouts. 

Wing  Stability.  Under  wing  sections,  the  subject  of  the 
center  of  pressure  movement  has  already  been  dealt  with. 
The  variation  of  the  center  of  pressure  with  the  angle 


332  STABILITY 

of  incidence  tends  to  destroy  longitudinal  stability  since 
the  center  of  pressure  does  not  at  all  times  pass  through 
the  center  of  gravity.  On  some  wings,  the  camber  is 
such  that  the  variation  in  the  position  of  the  center  of 


German    Stick  Control   With   Double   Grips.     A   Latch   on   the   Side   of  the 
Stick  Acts  on  a  Sector  So  That  the  Lever  Can  Be  Held  at  Any  Point. 
It  Is  Released  by  the  Pressure  of  the  Knees. 

pressure  is  very  little,  and  hence  these  are  known  as 
stable  wings.  A  reflex  curve  in  the  trailing  edge  of  a  wing 
reduces  the  center  of  pressure  movement,  and  swept  back 
wings  are  also  used  as  an  aid  in  securing  longitudinal 
stability.    Introducing  stagger  and  decalage  into  a  biplane 


STABILITY 


333 


pair  can  be  made  to  produce  almost  perfect  static  longi- 
tudinal stability.  It  should  be  noted  that  stability  ob- 
tained by  wing  and  camber  arrangements  is  static  only, 
and  requires  damping  surfaces  to  obtain  dynamic  stability. 
Manual  Controls.  In  flight,  the  aviator  has  three  con- 
trol surfaces  to  operate,  the  ailerons,  elevator,  and  rudder. 
In  the  usual  form  of  machine  the  ailerons  and  elevator  are 
operated  by  a  single  lever  or  control  column,  while  the 
rudder  is  connected  with  a  foot  bar.  In  the  smaller 
machines  "Stick  Control"  is  generally  used,  the  ailerons 


Form  of  Control  Used  on  the  Nieuport  Monoplane. 


and  elevator  being  moved  through  a  simple  lever  or  "Joy 
Stick"  which  is  pivoted  at  its  lower  end  to  the  floor.  The 
Deperdussin  or  *'Dep"  control  is  standard  with  the  larger 
machines  and  consists  of  an  inverted  **U"  form  yoke  on 
which  is  mounted  the  wheel  for  operating  the  ailerons. 
Stick  Control.  With  the  stick  pivoted  at  the  bottom, 
a  forward  movement  of  the  lever  causes  the  machine  to 
descend  while  a  backward  movement  or  pull  toward  the 
pilot  causes  the  aeroplane  to  head  up  or  ascend.  The  stick 
is  connected  with  the  elevators  with  crossed  wires,  so  that 


334 


STABILITY 


the  flaps  move  in  an  opposite  direction  to  the  "Stick." 
Moving  the  stick  from  side  to  side  operates  the  ailerons. 
Deperdussin  Control.  A  "U"  shaped  yoke,  either  of 
bent  wood  or  steel  tube,  is  pivoted  the  bearers  at  the 
sides  of  the  fuselage.  Wires  are  attached  to  the  bottom 
of  the  yoke  so  that  its  back  and  forth  movement  is  com- 
municated to  the  elevator  flaps.     On  the  top,  and  in  the 


-*^ 


left  s/de    Righf- s'id€ 
of  machine    of  tvach/nc 


Standard  "Dep"  Control  and  Movements  Used  in  the  U.  S.  A. 


■^^ 


sr/c/c 


Lcff  tide      RighfSidt 
of  machim     of  machine 


down 


down 


Standard  Stick  Control  and  Movements  Used  in  the  U.  S.  A. 


center  of  the  yoke,  is  pivoted  a  hand  wheel  of  the  auto- 
mobile steering  type.  This  is  provided  with  a  pulley  and 
is  connected  with  the  aileron  flaps  in  such  a  way  that 
turning  the  wheel  toward  the  high  wing  tip  causes  it  to 
descend.  Pushing  the  yoke  forward  and  away  from  the 
aviator  causes  the  machine  to  descend,  while  a  reverse 
movement  raises  the  nose.    The  "Dep"  control  is  reliable 


STABILITY  335 

and  powerful  but  is  bulky  and  heavy,  and  requires  a  wide 
body  in  order  to  allow  room  for  the  pilot. 

Rudder  Control.     Foot  bar  control  for  the  rudder  is 
standard  with  both  the  stick  and  Dep  controls.    The  foot 


Foot  Rudder  Bar  Used  in  the  Standard  H-3.      Courtesy   "Aerial  Age." 

bar  is  connected  with  the  rudder  in  such  a  way  that  the 
aeroplane  turns  opposite  to  the  movement  of  the  foot  bar 
in  the  manner  of  a  boat.  That  is,  pushing  the  right  end 
of  the  bar  forward  causes  the  machine  to  turn  toward 
the  right. 


CHAPTER  XVI. 
HEAD    RESISTANCE   CALCULATIONS. 

Effect  of  Resistance.  Resistance  to  the  forward  mo- 
tion of  an  aeroplane  can  be  divided  into  two  classes, 
(1)  The  resistance  or  drag  due  to  the  lift  of  the 
wings,  and  (2)  The  useless  or  "Parasitic"  resistance 
due  to  the  body,  chassis  and  other  structural  parts 
of  the  machine.  The  total  resistance  is  the  sum 
of  the  wing  drag  and  the  parasitic  resistance. 
Since  every  pound  of  resistance  calls  for  a  definite 
amount  of  power,  it  is  of  the  greatest  importance  to  re- 
duce this  loss  to  the  lowest  possible  amount.  The  adop- 
tion of  an  efficient  wing  section  means  little  if  there  is  a 
high  resistance  body  and  a  tangle  of  useless  struts  and 
wires  exposed  to  the  air  stream.  The  resistance  has  a 
much    greater    effect    on    the    power    than    the    weight. 

Weight  and  Resistance.  We  have  seen  that  the  aver- 
age modern  wing  section  will  lift  about  16  times  the 
value  of  the  horizontal  drag,  that  is,  an  addition  of  16 
pounds  will  be  equal  to  1  pound  of  head  resistance.  If, 
by  unnecessary  resistance,  we  should  increase  the  drag 
by  10  pounds,  we  might  as  well  gain  the  benefit  of  10  x 
16=160  pounds  of  useful  load.  The  higher  the  lift- 
drag  efficiency  of  the  wing,  the  greater  will  be  the  pro- 
portional loss  by  parasitic  resistance. 

Gliding  Angle.  The  gliding  angle,  or  the  inclination  of 
the  path  of  descent  when  the  machine  is  operating  without 
power,  is  determined  by  the  weight  and  the  total  head 
resistance.  With  a  constant  weight  the  angle  is  greatest 
when  the  resistance  is  highest.     Aside  from  considera- 

337 


I 


/ 


338  HEAD  RESISTANCE 

tions  of  power,  the  gliding  angle  is  of  the  greatest  impor- 
tance from  the  standpoint  of  safety.  The  less  the  resist- 
ance, and  the  flatter  the  angle  of  descent,  the  greater  the 
landing  radius. 

Numerically  this  angle  can  be  expressed  by:  Glide 
=  W/R,  where  W  =  the  weight  of  the  aeroplane,  and 
R  =  total  resistance.  Thus  if  the  weight  is  2500  pounds 
and  the  head  resistance  is  500  pounds,  the  rate  of  glide 
will  be:  2500/500^=5.  This  means  that  the  machine 
will  travel  forward  5  feet  for  every  foot  that  it  falls  ver- 
tically. If  the  resistance  could  be  decreased  to  100  pounds, 
the  rate  of  glide  would  be  extended  to  2500/100  =  25, 
or  the  aeroplane  would  travel  25  feet  horizontally  for 
every  foot  of  descent.  This  will  give  an  idea  as  to  the 
value  of  low  resistance. 

Resistance  and  Speed.     The  parasitic  resistance  of  a 

body  in  uniform  air  varies  as  the  square  of  the  velocity  at 

I  I  ordinary  flight  speeds.     Comparing  speeds  of  40  and  100 

[  miles  per  hour,  the  ratio  will  be  as  40-  is  to  100^=  1600: 

I  10,000^6.25,  that  is,  the  resistance  at  100  miles  per  hour 

Awill  be  6.25  times  as  great  as  at  40  miles  per  hour. 

The  above  remarks  apply  only  to  bodies  making  con- 
stant angle  with  the  air  stream.  Wings  and  lifting  sur- 
faces make  varying  angles  at  different  speeds  and  hence 
do  not  show  the  same  rate  of  increase.  In  carrying  a 
constant  load,  the  angle  of  the  aeroplane  wing  is  decreased 
as  the  speed  increases  and  up  to  a  certain  point  the 
resistance  actually  decreases  with  an  increase  in  the  speed. 
The  wing  resistance  is  greatest  at  extremely  low  speeds 
and  at  very  high  speeds.  As  the  total  resistance  is  made 
up  of  the  sum  of  the  wing  and  parasitic  resistance  at  the 
different  speeds,  it  does  not  vary  according  to  any  fixed 
law.  The  only  true  knowledge  of  the  conditions  existing 
through  the  range  of  flight  speeds  is  obtained  by  drawing 
a  curve  in  which  the  sums  of  the  drag  and  head  resistance 
are  taken  at  intervals. 


HEAD  RESISTANCE  339 

Resistance  and  Power.  The  power  consumed  in  over- 
coming parasitic  resistance  increases  at  a  higher  rate  than 
the  resistance,  or  as  the  cube  of  the  speed.  Thus  if  the 
speed  is  increased  from  40  to  100  miles  per  hour,  the  power 
will  be  increased  15.63  times.  This  can  be  shown  by  the 
following:  Let  V  ^  velocity  in  miles  per  hour,  H  = 
Horsepower,  K  =  Resistance  coefficient  of  a  body,  A  = 
Total  area  of  presentation,  and  R  =  resistance  in  pounds. 
Then  H  =  RV/375.  Since  R  =  KAV^  then  H  =  KAV^  x 
V/375  =  KAVV375. 

Resistance  and  Altitude.  The  resistance  decreases 
with  a  reduction  in  the  density  of  the  air  at  constant  speed. 
In  practice,  the  resistance  of  an  aeroplane  is  not  in  direct 
proportion  to  a  decrease  in  the  density  as  the  speed  must 
be  increased  at  high  altitudes  in  order  to  obtain  the  lift. 
The  following  example  given  by  Capt.  Green  will  show 
the  actual  relations. 

Taking  an  altitude  of  10,000  feet  above  sea  level  where 
the  density  is  0.74  of  that  at  sea  level,  the  resistance  at 
equal  speeds  will  be  practically  in  proportion  to  the  densi- 
ties. In  order  to  gain  sustentation  at  the  higher  altitude, 
the  speed  must  be  increased,  and  hence  the  true  resistance 
will  be  far  from  that  calculated  by  the  relative  densities. 
Assume  a  sea  level  speed  of  100  ft. /sec,  a  weight  of  3000 
pounds,  a  lift-drag  ratio  of  L/D  ^  15,  and  a  body  resist- 
ance of  40  pounds  at  sea  level. 

Because  of  the  change  in  density  at  10,000  feet,  the  fly- 
ing speed  will  be  increased  from  100  feet  per  second  to 
350  feet  per  second  in  order  to  obtain  sustentation.  With 
sea  level  density  this  increase  in  speed  (3.5  times)  would 
increase  the  body  resistance  3.5  x  3.5  =  12.25  times,  mak- 
ing the  total  resistance  12.25x40  =  490  pounds.  Since 
the  density  at  the  higher  altitude  is  only  0.74  of  that  at 
sea  level,  this  will  be  reduced  by  0.26,  or  0.26  x  490  =  364 
pounds.  Thus,  the  final  practical  result  is  that  the  sea 
level  resistance  of  the  body    (40  pounds)    is   increased 


340  HEAD  RESISTANCE 

9.1  times  because  of  the  speed  increase  necessary  for 
sustentation.  Since  the  wing  angle  and  hence  the  lift- 
drag  ratio  would  remain  constant  under  both  condi- 
tions, the  wing  drag  would  be  constant  at  both  altitudes, 
or  3000/15  =  200  pounds.  The  total  sea  level  resistance 
at  100  feet  per  second  is  200  -|-  40  =  240  pounds,  while  the 
total  resistance  at  10,000  feet  becomes  364  +  200  =  564 
pounds. 

The  speed  varies  as  the  square  root  of  the  change  in 
density  percentage.  If  V  =  velocity  at  sea  level,  v  = 
velocity  at  a  higher  level,  and  d  =  percentage  of  the  sea 

V 
level  density  at  the  higher  altitude,  then  v  =  — .    When 

Vd 
the  velocity  at  the  high  altitude  is  thus  determined,  the 
resistance  can  be  easily  obtained  by  the  method  given  in 
Capt.    Green's   article.     The   following  table   gives   the 
percentage  of  densities  referred  to  sea  level  density. 


Altitude 

Density 

Altitude 

Density 

Feet 

Percent 

Feet 

Percent 

Sea-level 

1.00 

7,500 

0.78 

1,000 

.97 

10,000 

.74 

2,000 

.95 

12,500 

.66 

3,000 

.91 

15,000 

.61 

5,000 

.85 

20,000 

.52 

If  the  velocity  at  sea  level  is  100  miles  per  hour,  the 
velocity  at  20,000  feet  will  be  100/0.72=139  miles  per 
hour,  where  0.72  is  the  square  root  of  the  density  per- 
centage, or  VO-52  =  .72  at  20,000  feet. 

Total  Parasitic  Resistance.  Aside  from  the  drag  of  the 
wings,  the  resistance  of  the  structural  parts,  body,  tail  and 
chassis  depends  upon  the  size  and  type  of  aeroplane.  A 
speed  scout  has  less  resistance  than  a  larger  machine 
because  of  the  small  amount  of  exposed  bracing,  although 
the  relative  resistance  of  the  body  is  much  greater.  The 
type  of  engine  also  has  a  great  influence  on  the  parasitic 
resistance.  The  following  gives  the  approximate  distribu- 
tion of  a  modern  fighting  aeroplane  : 


HEAD  RESISTANCE  3^1 

Body   62  percent 

Landing  gear 16 

Tail,  fin,  rudder 7         " 

Struts,  wires,  etc 15        '* 

The  body  resistance  is  by  far  the  greatest  item.  A 
great  part  of  the  body  resistance  can  be  attributed  to  the 
motor  cooling  system,  since  in  either  case  it  is  diverted 
from  the  true  streamline  form  in  order  to  accommodate 
the  radiator,  or  the  rotary  motor  cowl.  The  body  resist- 
ance is  also  influenced  by  the  necessity  of  accommodating 
a  given  cargo  or  passenger-carrying  capacity,  and  by  the 
distance  of  the  tail  surfaces  from  the  wings.  A  body  is 
not  a  streamline  form  when  its  length  greatly  exceeds  6 
diameters. 

Calculation  of  Total  Resistance.  The  nearest  approach 
that  we  can  make  to  the  actual  head  resistance  by  means 
of  a  formula  is  to  adopt  an  expression  in  the  form  of  R  = 
KV-  where  K  is  a  factor  depending  upon  the  size  and 
type  of  machine.  The  true  method  would  be  to  go  over 
the  planes  and  sum  up  the  individual  resistance  of  all  the 
exposed  parts.  The  parts  lying  in  the  propeller  slip  stream 
should  be  increased  by  the  increased  velocity  of  the  slip 
stream.  The  parasitic  resistance  of  biplanes  weighing 
about  1800  pounds  will  average  about,  R  =  0.036V^  where 
V  =  velocity  in  miles  per  hour.  Biplanes  averaging  2500 
pounds  give  R^0.048V-.  Machines  of  the  training  or 
2-seater  type  weigh  from  1800  to  2500  pounds,  and  have  an 
average  head  resistance  distribution  as  follows : 

Body,  radiators,  shields 35.5  percent. 

Tail  surface  and  bracing 14.9         *' 

Landing  gear 17.2 

Interplane  struts,  wires  and  fittings  23.6         " 
Ailerons,  aileron  bracing,  etc 8.8         " 

The  averages  in  the  above  table  differ  greatly  from  the 
values  given  for  the  high  speed  fighting  machine,  prin- 


342  HEAD  RESISTANCE 

cipally  because  of  the  large  control  surfaces  used  in  train- 

I  ing  machines,  and  the  difference  in  the  size  of  the  motors. 

'  With  the  wing  drag  being  equal  to  D  =  KxAV^  and 
the  total  parasitic  resistance  equal  to  R  =  KV^,  the  total 
resistance  can  be  expressed  by  Rt  =  KxAV^  +  KV-, 
where  K  ==  coefficient  of  parasitic  resistance  for  different 
types  and  sizes  of  machines.  The  value  of  K  for  training 
machines  will  average  0.036,  for  machines  weighing  about 
2500  pounds  K  =  0.048.  Scouts  and  small  machines  will 
be  safe  at  K  =  0.028.  The  wing  drag  coefficient  Kx  varies 
w4th  the  angle  of  incidence  and  hence  with  the  speed. 
For  example,  we  will  assume  that  the  wing  drag  (Kx)  of 
a  scout  biplane  at  100  miles  per  hour  is  0.00015,  that  the 
area  is  200  square  feet,  and  that  the  parasitic  resistance 
coefficient  is  K  =  0.028.  The  total  resistance  becomes: 
Rt  =  (0.00015  X  200  X  100  X  100)  +  0.028  x  100  x  100  =  300 
-\-  280  =  580  pounds.  The  formula  in  this  case  would  be 
Rt=^KxAV2  +  0.028V^ 

)  Strut  Resistance.  The  struts  are  of  as  nearly  stream- 
line form  as  possible.  In  practice  the  resistance  must  be 
compromised  with  strength,  and  for  this  reason  the  struts 
having  the  least  resistance  are  not  always  applicable  to 
the  practical  aeroplane.  From  the  best  results  published 
by  the  N.  P.  L.  the  resistance  was  about  12.8  pounds  per 
100  feet  strut  at  60  miles  per  hour.  The  width  of  the  strut 
is  1  inch.  A  rectangular  strut  under  the  same  conditions 
gave  a  resistance  of  104.4  pounds  per  100  feet.  A  safe 
value  would  be  25  pounds  per  100  feet  at  60  miles  per  hour. 

!  If  a  wider  strut  is  used,  the  resistance  must  be  increased 
in  proportion.  With  a  greater  speed,  the  resistance  must 
jbe  increased  in  proportion  to  the  squares  of  the  velocity, 
j  When  the  struts  are  inclined  with  the  wind,  the  resistance 
jis  much  decreased,  and  this  is  one  advantage  of  a  heavy 
stagger  in  a  biplane. 

The  "Fineness  ratio"  or  the  ratio  of  the  width  to  the 
depth  of  the  section  has  a  great  effect  on  the  resistance. 


HEAD  RESISTANCE  343 

With  the  depth  equal  to  twice  the  width  measured  across 
the  stream,  a  certain  strut  section  gave  a  resistance  of  24.8 
pounds  per  100  feet,  while  with  a  ratio  of  3.5  the  resistance 
was  reduced  to  11.4  pounds  per  hundred  feet.  Beyond 
this  ratio  the  change  is  not  as  great,  for  with  a  ratio  of 
4.6  the  resistance  only  dropped  to  11.2  pounds. 

Radiator  Resistance.  For  the  exact  calculation  of  the 
radiator  resistance  it  is  first  necessary  to  know  the  motor 
power  and  the  fuel  consumption  since  the  radiator  area, 
and  hence  the  resistance,  depends  upon  the  size  of  the 
motor  and  the  amount  of  heat  transmitted  to  the  jacket 
water.  An  aeronautic  motor  may  be  considered  to  lose 
as  much  through  the  water  jackets  as  is  developed  in  use- 
ful power,  so  that  on  this  basis  we  should  allow  about  1.6 
square  feet  of  radiation  surface  per  horsepower.  This 
figure  is  arrived  at  by  J.  C.  Hunsaker  and  assumes  that  the 
wind  speed  is  50  miles  per  hour  (73  feet  per  second). 
The  most  severe  cooling  condition  is  met  with  in  climbing 
at  low  speed,  and  it  is  here  assumed  that  50  miles  per  hour 
will  represent  the  lowest  speed  that  would  be  maintained 
for  any  length  of  time  with  the  motor  full  out.  For  a 
racing  aeroplane  that  will  not  climb  for  any  length  of 
time,  one-half  of  the  surface  given  above  will  be  sufiicient, 
and  if  the  radiator  is  placed  in  the  propeller  sHp  stream  it 
can  be  made  relatively  still  smaller  as  the  increased  pro- 
peller slip  at  rapid  rates  of  climb  partially  offsets  the 
additional  heating. 

In  the  above  calculations,  Hunsaker  does  not  take  any 
particular  type  of  radiator  into  consideration,  merely 
assuming  a  smooth  cooling  surface.  The  Rome-Turney 
Company  states  that  they  allow  1.08  square  feet  of  cooling 
surface  per  horsepower  for  honeycomb  radiators,  and  0.85 
square  feet  for  the  helical  tube  type.  The  surface  referred 
to  means  the  actual  suface  measured  all  over  the  tubes- 
and  cells,  and  does  not  refer  to  the  front  area  nor  the 
exterior  dimensions  of  the  radiator.     While  a  radiator 


344  HEAD  RESISTANCE 

may  be  made  25  percent  smaller  when  placed  in  the  slip- 
stream, the  resistance  is  increased  by  about  25  per  cent, 
with  a  very  small  saving  in  weight,  hence  the  total  saving 
is  small,  if  any.  Side  mounted  radiators  have  a  lower 
cooling  effect  per  square  foot  than  those  placed  in  any 
other  position,  owing  to  the  fact  that  the  air  must  pass 
through  a  greater  length  of  tube  than  where  the  broad 
side  faces  the  wind. 

In  the  radiator  section  tested  by  Hunsaker,  there  were 
about  64  square  feet  of  cooling  surface  per  square  foot 
of  front  face  area,  but  for  absolute  assurance  on  this  point 
one  should  determine  the  ratio  for  the  particular  type  of 
radiator  that  is  to  be  used.  The  Auto  Radiator  Manufac- 
turing Corporation,  makers  of  the  '*Flexo  copper  core 
radiators,  have  published  some  field  tests  made  under 
practical  conditions  and  for  different  types  and  methods 
of  mounting.  The  four  classes  of  radiators  described  are  : 
(1)  Front  Type,  in  which  the  radiator  is  mounted  in 
the  end  of  the  fuselage ;  (2)  Side  Type,  mounted  on  the 
sides  of  the  body ;  (3)  Overhead  Type,  mounted  above 
the  fuselage  and  near  the  top  plane ;  (4)  Over-Engine 
Type,  placed  above  and  connected  directly  to  the  motor, 
as  in  the  Standard  H-3. 

The  following  table  gives  the  eft'ectiveness  of  the  differ- 
ent mountings  in  terms  of  the  frontal  area  required  per 
horsepower  and  the  cooling  surface,  the  area  being  in 
square  inches  (Front  face  area  of  radiator).  Area  in  wind 
of  type  (3)  is  half  the  calculated  frontal  area  since  one 
core  lies  behind  the  other: 

Taking  the  value  of  the  Rome-Turney  honeycomb 
radiator  as  60  square  feet  of  cooling  surface  per  horse- 
power, the  frontal  area  per  horsepower  will  be  0.0169 
square  feet,  assuming  that  the  radiator  is  approximately 
6  inches  thick.  This  amounts  to  2.43  square  inches  of 
frontal  area  per  horsepower. 

Example.  Find  the  approximate  frontal  area  of  a  Rome- 


HEAD  RESISTANCE  345 

Turney  type  honeycomb  radiator  used  with  a  motor  giv- 
ing 100  brake-horsepower.  Find  Resistance  at  50  miles 
per  hour  {73  feet  per  second). 


Class   of   Mounting 

Square    Inches 
Per    H.P. 

Cooling  Surface 
Per  H.P.  Square  Inch 

Front   Type 

Side    Type 

Overhead     Type 

Over-Engine     Type 

4.00 
7.20 
2.70 
5.00 

117.00 
104.00 
112.00 
121.00 

Solution.  Area  =  A  =-  0.0169  HP  =  0.0169  x  100  =  1.69 
square  feet.  The  honeycomb  portion  of  surface  for  a 
square  radiator  of  the  above  area  will  measure  16.2"  x  16.2. 
Allowing  a  1-inch  water  passage  or  frame  all  around  the 
core,  the  side  of  the  completed  square  radiator  will  meas- 
ure 16.2"  -f  1"  -f  =  18.2". 

The  diameter  of  a  circular  radiator  core  of  the  same 

1.69x144 

area  will  be  17.4  inches,  since  D  =     .  Adding  the 

0.7854' 
water  passage,  the  overall  diameter  becomes  17.4  -f-  1  -|-  1 
=  19.4  inches.    The  round  honeycomb  front  radiator  used 
on  the  100  horsepower  Curtiss  Baby  Scout  measures  20 
inches. 

Hunaker  found  the  resistance  of  a  honeycomb  radiator 
to  be  R:=  0.000814  AV-,  there  being  4  honeycomb  cells 
per  square  inch.  A  =  area  of  radiator  in  square  feet,  and 
V  =  velocity  in  feet  per  second.  Adopting,  for  example, 
a  speed  of  73  feet  per  second,  and  an  area  equivalent  to  a 
19.4-inch  diameter  circular  radiator  as  above,  the  total 
resistance  becomes : 

R  =  0.000814  AV^  =  0.000814  x  3.1  x  {73  x  73)  =  13.32 
pounds,  at  50  M.  P.  H.  where  A  =  3.1  square  feet. 

Resistance  of  Chassis.  Disc  wheels  (Enclosed  spokes) 
have  a  resistance  of  about  one-half  that  of  open-wire 
wheels.    The  N.  P.  L.  and  Eiffel  have  agreed  that  the 


r 


346  HEAD  RESISTANCE 

resistance  of  a  wheel  approximating  26''  x  4"  has  a  resist- 
ance of  1.7  pounds  at  60  miles  per  hour  (Disc  type).  For 
any  other  speed,  the  wheel  resistance  will  be  R=1.7 
yV3600,  where  V  =  speed  in  miles  per  hour.  We  must 
also  take  into  consideration  the  axle,  chassis  struts,  wiring, 
/  ^  shock  absorbers,  etc.  The  itemization  of  the  chassis  re- 
sistance, as  given  by  the  N.  P.  L.  for  the  B.  E.-2  biplane  is 
as  follows  (60  miles  per  hour) : 

Wheels  2@1.75  pounds 3.5  pounds 

Axle 2.0      " 

Chassis  struts  and  connections 1.1       " 

Total  chassis  resistance(g60  MPH  .6.6  pounds 

At  any  other  speed,  the  resistance  for  the  complete 
chassis  can  be  given  by  the  formula  R  =  6.6VV3600. 
This  allowance  will  be  ample,  as  the  B.  E.-2  is  an  old  type 
and  is  equipped  with  skids. 

Interplane  Resistance.  The  interplane  struts  and  wires 
are  difficult  to  estimate  by  an  approximate  formula,  the 
only  exact  way  being  to  figure  up  each  item  separately 
from  a  preliminary  drawing.  The  resistance  varies  with 
the  form  of  the  strut  or  wire  section,  the  length,  and  the 
thickness.  The  fact  that  some  of  the  struts  lie  in  the  pro- 
peller slipstream,  and  some  outside  of  it,  makes  the  calcu- 
lation doubly  difficult.  The  only  recourse  that  we  have  at 
present  is  to  analyze  the  conditions  on  the  B.  E.-2.  With 
struts  approximating  true  streamline  form,  a  great  per- 
centage of  the  total  resistance  is  skin  friction,  and  as 
before  explained,  this  item  varies  at  a  lesser  rate  than 
the  square  of  the  speed. 

According  to  a  number  of  experiments  on  full  size 
biplanes  averaging  1900  pounds,  it  has  been  found  that  the 
interplane  resistance  (Struts,  wires  and  fittings)  amounts 
to  about  24  per  cent  of  the  total  parasitic  head  resistance 


HEAD  RESISTANCE 


347 


of  the  entire  machine,  the  drag  of  wings  not  being  in- 
cluded. The  maximum  observed  gave  29  per  cent  and  the 
minimum  15  per  cent.  The  resistance  of  the  interplane 
bracing  of  speed  scouts  will  be  considerably  less  in  pro- 
portion, as  there  are  fewer  exposed  struts  and  cables  on 
this  type,  the  resistance  probably  averaging  15  per  cent  of 
the  total  head  resistance.    Based  on  these  figures  the  re- 

INTERPLANE  RESISTANCE  OF  BIPLANE  B.  E.-2 
AT  60  M.  P.  H. 


Location  of  Items 

Length  of 
Strut  in  Ft. 

Total   Running 
Length 

Resistance 
in    Lbs. 

Total    Resist- 
ance of  Group 

In  translation  stream. 
In  translation  stream. 
In  translation  stream. 

8@6'0"=- 
4@4'0"= 
6@3'0"= 

48'0" 
16'0" 
18'0" 

4.20 
1.40 
1.60 

In  slipstream  only... 
In  slipstream  only... 

Length 
4@4'0"= 
4@3'0"= 

=  82'0" 
16'0" 
12'0" 

7.20 
2.30 
1.31 

7.20  lbs. 

Length 

=  28'0" 

3.61 

3.61  lbs. 

sistance  of  the  interplane  bracing  can  be  expressed  by  the 
following  formula,  in  which  I  =  resistance  of  interplane 
bracing  in  pounds,  and  V  =  translational  speed  in  miles 
per  hour: 

I  =  0.009V'  (For  two-place  biplanes  weighing  1900 
pounds). 

I  ^  0.0054V"  (For  biplane  speed  scouts  or  racing  type 
biplanes). 

Strut  Resistance.  The  above  estimate  includes  wiring, 
strut  fittings,  etc.,  complete,  and  also  takes  the  effect  of 
the  slipstream  into  consideration.  A  more  accurate  esti- 
mate can  be  made  on  the  basis  of  strut  length.  To  obtain 
this  unit  value  we  have  recourse  to  the  B.  E.-2  tests.  The 
translational  speed  in  60  miles  per  hour  (88  feet  per  sec- 
ond) and  the  slipstream  is  taken  at  25  feet  per  second. 
This  gives  a  total  velocity  in  the  slipstream  of  113  feet  per 
second.  The  struts  are  \%  inches  wide,  and  vary  in  length 
from  3'  — (y^  to  6'  —  0".    In  the  slipstream  the  increased 


348  HEAD  RESISTANCE 

velocity  increases  the  resistance  of  the  items  by  64  per 
cent. 

Total  running  length  =110'  —  0''.  Total  resistance  = 
10.81  pounds.  The  resistance  per  foot  =  10.81/110  = 
0.099  pounds. 

Resistance  of  Wire  and  Cable.  In  this  estimate  we  will 
take  the  resistance  given  in  the  B.  E.-2  tests,  since  values 
are  given  in  the  slipstream  as  well  as  for  the  outer  por- 
tions. In  the  translational  stream  there  is  240'  —  0"  of 
cable,  70' =  0"  of  No.  12  solid  wire,  and  52  turnbuckles, 
the  total  giving  a  resistance  of  38.10  pounds.  In  the  slip- 
stream there  is  50'  —  0"  of  cable  and  30'  —  0"  of  solid 
wire  with  a  resistance  of  11.00  pounds.  The  total  wire 
and  cable  resistance  for  the  wings  is  therefore  49.10 
pounds.  The  resistance  of  the  wire  and  cable  combined  is 
0.127  pounds  per  running  foot. 

Summary  of  Interplane  Resistance.  The  total  inter- 
plane  resistance  includes  the  struts,  wires,  cables  and  turn- 
buckles,  a  portion  of  which  are  m  the  slipstream.  Since 
the  total  head  resistance  of  the  entire  machine  (B.  E.-2) 
is  140  pounds  at  60  M.  P.  H.,  and  the  interplane  resistance 
=  10.81  +  49.10=  59.91  pounds,  the  relation  of  the  inter- 
plane resistance  to  the  total  resistance  is  43  per  cent.  This 
is  much  higher  than  the  average  (24  per  cent),  but  the 
B.  E. — 2  is  an  old  type  of  machine  and  the  number  of 
struts  and  wires  were  much  greater  than  with  modern 
aeroplanes. 

Control  Surface  Resistance.  The  resistance  of  the  con- 
trol surfaces  is  a  variable  quantity,  since  so  much  depends 
upon  the  arrangement  and  form.  Another  variation  oc- 
curring among  machines  of  the  same  make  and  type  is  due 
to  the  various  angles  of  the  surfaces  during  flight,  or  at 
least  during  the  time  that  they  are  used  in  correcting  the 
attitude  of  the  machine.  With  the  elevator  flaps  or 
ailerons  depressed  to  their  fullest  extent,  the  drag  is  many 
times  that  with  the  surfaces  in  ''neutral,"  and  as  a  general 


HEAD  RESISTANXE  349 

thing  the  controls  are  depressed  at  the  time  when  the 
power  demand  is  the  greatest — that  is,  on  landing,  flying 
slow,  or  in  ''getting  off." 

Ailerons  ''in  neutral"  can  be  considered  as  being  an 
integral  part  of  the  wings  when  they  are  hinged  to  the 
wing  spar.  In  the  older  types  of  Curtiss  machines  the 
ailerons  were  hinged  midway  between  the  planes  and  the 
resistance  was  always  in  existence,  whether  the  ailerons 
were  in  neutral  or  not.  Wing  warping,  in  general,  can  be 
assumed  as  in  the  case  where  the  wings  and  ailerons  are 
combined.  With  ailerons  built  into  the  wings,  the  re- 
sistance of  the  ailerons,  and  their  wires  and  fittings,  can 
be  taken  as  being  about  4  per  cent  of  the  total  head  resist- 
ance. With  the  aileron  located  between  the  two  wings, 
the  resistance  may  run  as  high  as  20  per  cent  of  the 
total. 

Like  the  ailerons,  the  elevator  surfaces  and  rudder  are 
variable  in  attitude  and  therefore  give  a  varying  resist- 
ance. In  neutral  attitude  the  complete  tail,  consisting  of 
the  rudder,  stabilizer,  elevator,  fin  and  bracing,  will  aver- 
age about  15  per  cent  of  the  total  resistance,  it  being 
understood  that  a  non-lifting  stabilizer  is  fitted.  With 
lifting  tails  the  resistance  will  be  increased  in  proportion 
to  the  load  carried  by  the  stabilizer.  In  regard  to  the  tail 
resistance  it  should  be  noted  that  these  surfaces  are  in  the 
slipstream  and  are  calculated  accordingly,  although  the 
velocity  of  the  slipstream  is  somewhat  reduced  at  the 
point  where  it  encounters  the  tail  surfaces.  The  total  tail 
resistance  of  the  B.  E.-2  is  given  as  3.3  pounds. 

Resistance  of  Seaplane  Floats.  The  usual  type  of  sea- 
plane with  double  floats  may  be  considered  as  having 
about  12  per  cent  higher  resistance  than  a  similar  land 
machine.  Some  forms  of  floats  have  less  resistance  than 
others,  owing  to  their  better  streamline  form,  but  the 
above  figure  will  be  on  the  safe  side  for  the  average  pon- 
toon.   Basing  our  formula  on  a  12  per  cent  increase  on  the 


350  HEAD  RESISTANCE 

total  head  resistance,  the  formula  for  the  floats  and  bracing 
will  become: 

Rf  =  0.00436V2  where  Rf  =  resistance  of  floats  and  fit- 
tings. 

Body  Resistance.  This  item  is  probably  the  most  diffi- 
cult of  any  to  compute,  owing  to  the  great  variety  of 
forms,  the  difference  in  the  engine  mounting,  and  the  dis- 
position of  the  fittings  and  connections.  The  resistance 
of  the  pilot's  and  passenger's  heads,  wind  shields,  and  pro- 
peller arrangement  all  tend  to  increase  the  difficulty  of 
obtaining  a  correct  value.  Aeroplanes  with  rotary  air- 
cooled  motors,  or  with  large  front  radiators  have  a  higher 
resistance  than  those  arranged  with  other  types  of  motors 
or  radiator  arrangements.  Probably  the  item  having  the 
greatest  influence  on  the  resistance  of  the  fuselage  is  the 
ratio  of  the  length  to  the  depth,  or  the  "fineness  ratio." 
In  tractor  monoplanes  and  biplanes,  of  the  single  pro- 
peller type,  the  body  is  in  the  sHpstream,  and  compensa- 
tion must  be  made  for  this  factor. 

If  it  were  not  for  the  motor  and  radiator,  the  tractor 
fuselage  could  be  made  in  true  dirigible  streamline  form, 
and  would  therefore  present  less  resistance  than  the  pres- 
ent forms  of  "practical"  bodies.  The  necessity  of  placing 
the  tail  surfaces  at  a  fixed  distance  from  the  wings  also 
involves  the  use  of  a  body  that  is  longer  in  proportion  than 
a  true  streamline  form,  and  this  factor  alone  introduces 
an  excessive  head  resistance.  The  ideal  ratio  of  depth 
to  length  would  seem  to  range  from  1  to  5.5  or  1  to  6.  The 
fineness  ratio  of  the  average  two-seat  tractor  is  considera- 
bly greater  than  this,  ranging  from  1  to  7.5  or  8.5.  A 
single-seat  machine  of  the  speed-scout  type  can  be  made 
much  shorter  and  has  more  nearly  the  ideal  proportions. 

The  only  possible  way  of  disposing  of  this  problem  is 
to  compare  the  results  of  wind  tunnel  tests  made  on  dif- 
ferent types  of  bodies,  and  even  with  this  data  at  hand 


HEAD  RESISTANCE  351 

a  liberal  allowance  should  be  made  because  of  the  in- 
fluence of  the  connections  and  other  accessories.  Eiffel, 
the  N.  P.  L.,  and  the  Massachusetts  Institute  of  Tech- 
nology have  made  a  number  of  experiments  with  scale 
models  of  existing  aeroplane  bodies.  It  is  from  these 
tests  that  we  must  estimate  our  body  resistance,  hence  a 
table  of  the  results  is  attached,  the  approximate  outlines 
being  shown  by  the  figures. 

As  in  calculating  the  resistance  of  other  parts,  the  re- 
sistance of  the  body  can  be  expressed  by  R  =  KxAV-, 
where  Kx  =  coefficient  of  the  body  form,  A  =  Cross-sec- 
tional area  of  body  in  square  feet  (Area  of  presentation), 
and  V  =  velocity  in  miles  per  hour.  The  area  A  is  ob- 
tained by  multiplying  the  body  depth  by  the  width.  The 
"area  of  presentation"  of  a  body  2'  —  6"  wide  and  3'  —  0" 
deep  will  be  2.5  x  3  =  7.5  square  feet. 

The  experimental  data  does  not  give  a  very  ready  com- 
parison between  the  different  types,  as  the  bodies  not  only 
vary  in  shape  and  size,  but  are  also  shown  with  different 
equipment.  Some  have  tail  planes  and  some  have  not; 
two  are  shown  with  the  heads  of  the  pilot  and  passenger 
projecting  above  the  fuselage,  while  the  remainder  have 
either  a  simple  cock-pit  opening  or  are  entirely  closed. 
The  presence  of  the  propeller  in  two  cases  may  have  a 
great  deal  to  do  with  raising  the  value  of  the  experimental 
results.  The  propeller  was  stationary  during  the  tests, 
but  it  was  noted  that  the  resistance  was  considerably  less 
when  the  propeller  was  allowed  to  run  as  a  windmill, 
driving  the  motor.  This  latter  condition  would  correspond 
to  the  resistance  in  gliding  with  the  motor  cut  off.  In 
all  cases,  except  the  Deperdussin,  the  bodies  are  covered 
with  fabric,  and  the  sagging  of  the  cloth  in  flight  will 
probably  result  in  higher  resistance  than  would  be  indi- 
cated by  the  solid  wood  or  metal  model  used  in  the  tests. 
The  pusher  type  bodies  give  less  resistance  than  the 
tractors,  but  the  additional  resistance  of  the  outriggers 


352 


HEAD  RESISTANCE 


and  tail  bracing  will  probably  bring  the  total  far  above  the 
tractor  body. 

In  the  accompanying  body  chart  are  shown  7  repre- 
sentative bodies :  (a)  Deperdussin  Monocoque  Mono- 
plane Body,  a  single-seater ;  (b)  N.  P.  L.-5  Tractor  Biplane 
Body,  single-seater ;  (c)  B.  F.-36  Dirigible  Form,  without 
propeller  or  cock-pit  openings ;  (d)  B.  E.-3  Two-Place 
Tractor  Body,  with  passenger  and  pilot;  (e)  Curtiss  JN 
Type  Tractor  Body,  with  passengers,  chassis  and  tail ;  (f ) 
Farman  Pusher  type,  with  motor,  propeller  and  exposed 
passengers;  (g)  N.  P.  L.  Pusher  Body,  bare.  Body  (a) 
was  tested  with  a  1/5  scale  model  at  a  wind  tunnel  speed  of 
2S  meters  per  second,  the  resistance  of  the  model  being 
0.377  kilograms  (0.83  pounds).  Body  (d)  in  model  form 
was  1/16  scale  and  was  tested  at  20.5  miles  per  hour,  at 
which  speed  the  resistance  was  0.0165  pounds.  Model  (e) 
was  1/12  scale  and  was  tested  at  30  miles  per  hour.  These 
varying  test  speeds,  it  will  be  seen,  do  not  allow  of  a  very 
accurate  means  of  comparison.  The  resistance  of  model 
(e)  was  0.1365  pounds  at  the  specified  wind-tunnel  air 
speed. 

TABLE  OF  BODY  RESISTANCE. 


Full    Size    Re- 

Coefficient  of 

Fineness 

Model 

Tests 

sistance 

l-'ig. 

Name  of  Body 

Resistance    Kx 

Ratio 

Scale 

Speed 
M.P.H. 

A  =  8    Sq.    Ft. 
@   100  M.P.H. 

(a) 

Deperdussin     .... 

0.000541 

5.60 

1/5 

62.6 

52.16  lbs. 

(b) 

N.P.L.-5    

0.000444 

5.50 

33.24  lbs. 

(c) 

B.F.-36  Dirig 

0.000258 

5.75 

20.50  lbs. 

(H) 

B.E.-3    

0.000720 

7.35 

1/16 

20.5 

97.60  lbs. 

(e) 

Curtiss  Type  JN. 

0.002730 

8.20 

1/12 

30.0 

218.40  lbs. 

(t) 

Farman    Pusher.  . 

0.000855 

3.20 

67.59  lbs. 

(g) 

N.P.L.    Pusher... 

0.000271 

3.00 

21.60  lbs. 

The  speeds  given  in  the  above  table  are  simply  transla- 
tional  speeds,  and  are  not  corrected  for  slipstream  velocity. 
With  a  slipstream  of  25  per  cent,  increase  the  body  resist- 
ance by  40  per  cent.  It  would  be  safe  to  add  an  addi- 
tional 10  per  cent  to  make  up  for  projecting  fittings,  baggy 
fabric,  and  scale  variations. 


HEAD  RESISTANCE 


353 


Since  a  body  of  approximately  streamline  form  has  a 
considerable  percentage  of  skin  friction,  scale  corrections 
for  size  and  velocity  are  even  of  more  importance  than 
with  wing  sections.  No  wind-tunnel  experiments  can  de- 
termine the  resistance  exactly  because  of  the  uncertainty 
of  the  scale  factor.  The  resistance  as  given  in  the  table  is 
also  affected  by  the  proximity  of  the  wing  and  tail  sur- 


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A- CROSS-SECTIONAL  AREA  /N  S<p.  FECT. 


Chart  Showing  Forms  of  7  Typical  Aeroplane  Fuselage. 

faces,  and  by  projections  emanating  from  the  motor  com- 
partment. It  will  be  noted  that  the  dirigible  form  B.  F.-36 
is  markedly  better  than  any  of  the  others,  being  almost  of 
perfect  streamline  form.  The  nearest  approximation  to 
the  ideal  form  is  N.  P.  L.-5,  which  has  easy  curves,  low 


354  HEAD  RESISTANCE 

resistance,  and  is  fairly  symmetrical  about  the  center  line. 
Because  of  their  small  size,  the  pusher  bodies  or  "nacelles" 
have  a  small  total  resistance,  but  the  value  of  Kx  is  high. 
Problem.  Find  the  resistance  of  a  Curtiss  Tractor  Type 
JN  body  with  a  breadth  of  2'  —  6'  and  a  depth  of  3'  —  3', 
the  speed  being  90  miles  per  hour.  The  slipstream  is  as- 
sumed to  be  25  per  cent,  with  an  additional  10  per  cent 
for  added  fabric  loss,  etc. 


Typical  Stream  Line  Strut  Construction. 

Solution.  The  cross-sectional  area  =  2'  —  6"  x  3'  —  3" 
=  A  =  8.13  square  feet.  The  velocity  of  translation  is  90 
M.  P.  H.,  or  V^  =  8100.  The  value  of  the  resistance  coeffi- 
cient is  taken  from  the  table,  Kx^O.00273.  The  total  resist- 
ance R  ==  KxAV^  =  0.00273  x  8.13  x  8100  =  178.2  pounds. 
Since  a  slipstream  of  25  per  cent  increases  the  resistance 
by  40  per  cent,  the  resistance  in  the  slipstream  is  178.2  x 
1.4  =  249.48  pounds.  The  addition  of  the  10  per  cent  for 
extra  friction  makes  the  total  resistance  =  249.48  x  1.1  = 
274.43  pounds.  The  resistance  of  this  body,  used  with 
"twin"  motors,  would  be  178.2x1.1  =  196.02,  but  as  a 
tractor  with  the  body  in  the  sHpstream,  the  resistance 
would  be  equal  to  274.43  pounds  as  calculated  above. 


CHAPTER  XVII. 
POWER  CALCULATIONS. 

Power  Units.  Power  is  the  rate  of  doing  work.  If  a 
force  of  10  pounds  is  applied  to  a  body  moving  at  the  rate 
of  300  feet  per  minute,  the  power  will  be  expressed  by 
10x300  =  3000  foot-pounds  per  minute.  As  the  figures 
obtained  by  the  foot  and  pound  units  are  usually  incon- 
veniently large,  the  "Horsepower"  unit  has  been  adopted. 
A  horsepower  is  a  unit  that  represents  work  done  at  the 
rate  of  33,000  foot-pounds  per  minute,  or  550  foot-pounds 
per  second.  Thus  if  a  certain  aeroplane  offers  a  resist- 
ance of  200  pounds,  and  flies  at  the  rate  of  6000  feet  per 
minute,  then  the  work  done  per  minute  will  be  equal  to 
200  X  6000=  1,200,000  foot-pounds.  Since  there  are  33.000 
foot-pounds  of  work  per  minute  for  each  horsepower,  the 
horsepower  will  be  :  1200000/33000  =  36.3. 

As  aeroplane  speeds  are  usually  given  in  terms  of  miles 
per  hour,  it  will  be  convenient  to  convert  the  foot-minute 
unit  into  the  mile  per  hour  unit.  If  H  ==  horsepower,  R  =: 
resistance  of  aeroplane,  and  V  =  miles  per  hour,  then 
H  =  RV/375.  the  theoretical  horsepower,  without  loss. 
If  an  aeroplane  flies  at  100  miles  per  hour  and  requires  a 
propeller  thrust  of  300  pounds,  then  the  horsepower  be- 
comes : 

H  =  RV/375  =  300x100/375  =  80  horsepower.  This 
is  the  actual  power  required  to  drive  the  machine,  but  is 
not  the  engine  power,  as  the  engine  must  also  supply  the 
losses  due  to  the  propeller.  The  propeller  losses  are  gen- 
erally expressed  as  a  percentage  of  the  total  power 
supplied.  The  percentage  of  useful  power  is  known  as 
the  "Efficiency." 

355 


356  POWER 

The  efficiency  of  the  average  aeroplane  propeller  will 
vary  from  0.70  to  0.80.  If  e  =  propeller  efficiency  ex- 
pressed as  a  decimal,  the  motor  horsepower  becomes : 
H==  RV/375e.  To  obtain  the  motor  horsepower,  divide 
the  theoretical  horsepower  by  the  efficiency.  Using  the 
complete  formula  for  the  solution  of  an  example  in  which 
the  flight  speed  is  100  M.  P.  H.,  the  resistance  ?25  pounds 
and  the  efficiency  0.75,  we  have : 
I  /      H  =  RV/375e  =  225  x  100/375  x  0.75  =  80  horsepower. 

Power  Distribution.  Since  power  depends  upon  the 
total  resistance  to  be  overcome,  part  of  the  power  will  be 
used  for  driving  the  lifting  surfaces  and  a  part  for  over- 
coming the  parasitic  resistance.  The  power  required  for 
driving  the  wings  depends  upon  the  angle  of  incidence, 
since  the  drag  varies  with  every  angle.  The  wing  power 
varies  with  every  flight  speed,  owing  to  the  changes 
in  angle  made  necessary  to  support  the  constant  load. 
The  power  for  the  wings  will  be  least  at  the  speed  and 
angle  that  corresponds  to  the  greatest  lift-drag  ratio. 
Owing  to  the  low  value  of  the  L/D  at  very  small  and  very 
large  angles,  the  power  requirements  will  be  excessive  at 
extremely  low  and  high  speeds. 

As  the  parasitic  resistance  increases  as  the  square  of 
the  speed,  the  power  for  overcoming  this  resistance  will 
vary  as  the  cube  of  the  speed.  It  is  the  parasitic  resistance 
that  really  limits  the  higher  speeds  of  the  aeroplane,  since 
it  increases  very  rapidly  at  velocities  of  over  60  miles  per 
hour. 

The  total  power  at  any  speed  is  the  sum  of  the  wing 
power  and  power  required  for  the  parasitic  resistance. 
Owing  to  variations  in  the  wing  drag  and  resistance  at 
every  point  within  the  flight  range,  it  is  exceedingly  diffi- 
cult to  directly  calculate  the  total  power  at  any  particular 
speed.  The  wing  drag  and  the  resistance  should  be  cal- 
culated for  every  speed,  and  then  laid  out  by  a  graph  or 
curve.    The  minimum  propeller  thrust,  or  the  minimum 


POWER 


35i 


total  resistance,  occurs  approximately  at  the  speed  where 
the  body  resistance  and  wing  drag  are  equal.  The 
minimum  horsepower  occurs  at  a  low  speed,  but  not  the 
lowest  speed,  and  this  will  differ  with  every  machine. 

Fig.  1  is  a  set  of  performance  curves  drawn  from  the 
results  of  tests  on  a  full  size  Bleriot  monoplane.    At  the 


Fig.  1.  Power  Chart  of  Bleriot  Monoplane,  With  Outline  of  Wing  Section. 
The  Results  Were  Taken  From  Full  Size  Tests  Made  by  the  English 
Government. 

bottom  the  horizontal  row  of  figures  gives  the  horizontal 
speed  in  feet  per  second.  The  first  column  to  the  left  is 
the  horsepower,  and  the  second  column  is  the  resistance 
or  drag  in  pounds.  The  four  curves  represent  respective- 
ly the  body  resistance,  wing  or  "plane"  resistance,  horse- 
power, and  total  resistance.  The  horizontal  line  "AV" 
shows  the  available  horsepower.  It  will  be  noted  that  the 
body  resistance  increases  steadily  from  9  pounds  at  50  feet 


358  POWER 

per  second,  to  180  pounds  at  100  feet  per  second.  The 
wing  resistance,  on  the  other  hand,  decreases  from  350 
pounds  at  56  feet  per  second  to  a  minimum  of  130  pounds 
at  83  feet  per  second.  It  will  be  noted  that  the  angles  of 
incidence  are  marked  along  the  wing-drag  curve  by  small 
circles.  The  incidence  is  6°  at  75  feet  per  second,  and  4° 
at  a  little  less  than  85  feet  per  second. 

The  available  horsepower  "AV"  is  42.  This  is  shown 
as  a  straight  line,  although  in  the  majority  of  cases  it  is 
slightly  curved  owing  to  variations  in  power  at  the  higher 
speeds,  and  to  variations  in  the  propeller  efficiency.  At  90 
feet  per  second  the  actual  horsepower  curve  crosses  the 
line  of  available  horsepower  "AV."  Beyond  this  point 
horizontal  flight  is  no  longer  possible,  as  the  power  re- 
quirements would  exceed  the  available  horsepower.  It 
will  be  noted  that  the  lowest  total  resistance  occurs  near 
the  point  where  the  body  and  wing  resistance  curves  in- 
tersect, or  in  other  words,  where  the  body  and  wing  resist- 
ance are  equal.  The  minimum  horsepower  takes  place  at 
63  feet  per  second,  or  at  a  point  nearly  1/3  between  the 
lowest  flight  speed  and  the  highest  speed  attained  by  the 
available  horsepower  in  horizontal  flight  (90  ft/sec). 

The  actual  range  of  flight  speeds  is  limited  to  points 
between  the  intersection  of  the  "Horsepower  required" 
curve,  and  the  "Available  horsepower"  curve.  By  increas- 
ing the  propeller  efficiency,  or  by  increasing  the  power 
of  the  motor,  the  available  horsepower  line  is  raised  and 
the  flight  range  increased. 

Horsepower  For  Climbing.  Up  to  the  present  we  have 
only  considered  horizontal  flight.  The  power  available 
for  climbing  is  the  difference  between  the  power  required 
to  maintain  horizontal  flight  at  any  speed,  and  the  actual 
horsepower  that  can  be  delivered  by  the  propeller.  Thus, 
if  the  actual  power  delivered  by  a  motor  through  the  pro- 
peller is  85  horsepower,  and  the  power  required  for  hori- 
zontal flight  at  that  speed  is  45,  then  we  have :  85  —  45  = 


POWER  359 

40  horsepower  available  for  climbing.  Since  the  difference 
between  the  driving  power  and  the  power  required  for 
horizontal  flight  is  less  at  extremely  low  and  high  speeds, 
it  is  evident  that  we  will  have  a  minimum  climbing  re- 
serve at  the  high  and  low  speeds.  Consulting  the  powei 
curve  for  the  Bleriot  monoplane,  we  see  that  the  power 
required  at  56  feet  per  second  is  40  horsepower,  and  at 
85  feet  per  second  it  is  38  horsepower.  At  the  low  speed 
we  have  a  climbing  reserve  of  44  —  40  ==4  H.  P.,  and  at 
the  higher  speed  44  —  38  =  6  H.  P.  The  maximum  availa- 
ble horsepower  "AV"  is  44  horsepower.  The  minimum 
horizontal  power  required  is  found  at  63  feet  per  second, 
the  climbing  reserve  at  this  point  being  44  —  28=16 
horsepower.  At  55  feet  per  second,  and  at  90,  we  would 
not  be  able  to  climb,  as  we  would  only  have  sufficient 
power  to  maintain  horizontal  flight. 

If  W  =  total  weight  of  aeroplane,  c  =  climbing  speed, 
and  H  =  horsepower  reserve  for  climbing,  then  the  climb- 
ing speed  with  a  constant  air  density  will  be  expressed  by : 
c  =  33000H/\V.  Assuming  that  the  weight  of  the  Bleriot 
monoplane  is  800  pounds,  and  that  we  are  to  climb  at  the 
speed  of  the  greatest  power  reserve  (16  horsepower),  our 
rate  of  climb  is: 

c  =  33000H/W  =  33000  x  16/800  =  660  feet  per  minute. 
It  should  be  understood  that  this  is  the  velocity  at  the 
beginning  of  the  climb.  After  prolonged  climbing  the  rate 
falls  off  because  of  diminishing  power  and  increasing 
speed.  Much  depends  upon  the  engine  performance  at 
the  higher  altitudes,  so  that  the  reserve  power  for  climb 
usually  diminishes  as  the  machine  rises,  and  hence  the 
rate  of  climb  diminishes  in  proportion. 

The  following  table  taken  from  actual  flying  tests  will 
show  how  the  rate  of  climb  decreases  with  the  altitude. 
These  machines  were  equipped  with  150  H.  P.  Hispano- 
Suisa  motors : 

It  will  be  noted  that  the  S.  P.  A.  D.  and  the  Bleriot  hold 


360  POWER 

their  rate  of  climb  constant  up  to  7800  feet  altitude,  which 
is  a  feat  that  is  undoubtedly  performed  by  varying  the 
compression  of  the  engine. 

Besides  increasing  the  power,  the  rate  of  climb  can  also 
be  increased  by  decreasing  the  weight  of  the  aeroplane. 

Time  Required  for  Alti-     Rate  of  Climb  in  Feet 
tude    of  Per  Minute 

Name   of  Horizontal  o  o  o  o  o  o 

Machine  Speed  °  §  S  §  S  ^ 

ro  tC  .-T  jC  ^*  <^ 

Nieuport  97  M.P.H.  4'45"  12'20"  25'03"  820  630  465 

*F.  B.  A 89  M.  P.H.  6'40"  14'50"  22'30"  585  525  520 

B.  E 99  M.P.H.  3'50"  9'04"  16'00"  1,025  860  730 

S.  P.  A.  D 124  M.P.H.  3'00"  6'00"  10'30"  1,300  1,300  1,110 

Caudron  88  M.P.H.  7'10"  17'00"  28'40"  540  460  409 

Bleriot  124  M.P.H.  3'00"  6'00"  10'30"  1,300  1,300  1,110 

Maximum  Altitude.  The  maximum  altitude  to  which  a 
machine  can  ascend  is  known  as  its  "Ceiling."  This  again 
depends  on  both  the  aeroplane  and  the  motor,  but  prin- 
cipally on  the  latter.  It  has  been  noted  that  machines 
having  the  greatest  rate  of  climb  also  have  the  greatest 
ceiling.  Thus  the  ceiling  of  a  fast  climbing  scout  is  higher 
than  that  of  a  larger  and  slower  machine.  Based  on  this 
principle,  a  writer  in  "Flight"  has  developed  the  follow- 
ing equation  for  ceiling,  which,  of  course,  assumes  a  uni- 
form decrease  in  density.  Let  H  =  maximum  altitude, 
h  =  the  altitude  at  any  time  t  after  the  start  of  the  climb, 
and  a  =  the  altitude  after  a  time  equal  to  twice  the  time 
t;  then : 

H  =      h       Approximate  values  of  h  and  a  may  be  had 


2  — a/h 
from  the  following  table,  which  are  the  results  of  a  test  on 
a  certain  aeroplane : 

Time  (Minutes) 
0.0    2.5    5.0    7.5     10.0    12.5    15.0    17.5    20.0    22.5 

Altitude  (Feet) 
0.0  3300  6150  8730  10760  12610  14190  15530  16650  17600 


POWER  361 

If  we  assume  that  the  height  is  10760  feet  after  the  first 
10  minutes,  and  that  the  altitude  after  twice  this  time  (20 
minutes)  is  16650  feet,  then  the  maximum  ceiling  attained 
will  be : 

H  =       h       =  10760  =  23,770  feet.  The  use 


2  — a/h      2—16650/10760 

of  this  formula  requires  that  the  climb  be  known  for  cer- 
tain time  intervals  before  the  ceiling. 

Gliding  Angle.  The  gliding  angle  of  the  wings  alone  is 
equal  to  the  lift-drag  ratio  at  the  given  angle.  The  best 
or  ''Flattest  gliding  angle"  is,  of  course,  the  best  lift-drag 
ratio  of  the  wing  —  say  on  the  average  about  1  in  16.  The 
gliding  angle  of  the  complete  machine  is  considerably  less 
than  this,  owing  to  the  resistance  of  the  body  and  struc- 
tural parts.  This  generally  reduces  the  actual  angle  to 
less  than  12,  and  in  most  cases  between  6  and  8.  Ex- 
pressed in  terms  of  degrees,  tan  0  =  R/W  where  R  =  head 
resistance  and  W  =  weight  in  pounds. 

Fig.  2  is  a  diagram  giving  the  gliding  force  diagram. 
The  plane  descends  along  the  gliding  path  AC,  making  the 
angle  of  incidence  (0).  When  in  horizontal  flight,  the  lift 
is  along  OL  and  the  weight  is  0\V.  When  descending 
on  the  gliding  path  the  lift  maintains  the  same  relation 
with  the  wing,  but  the  relative  angle  of  the  weight  is 
altered.  The  weight  now  acts  along  OG.  The  drag  is  rep- 
resented by  OD,  with  the  propeller  thrust  OP  equal  and 
opposite  to  it.  With  the  weight  constant,  the  lift  OL  is  de- 
creased by  the  angle  so  that  the  total  life  ==  L  =  W  cos  ©. 
The  action  of  the  weight  W  produces  the  propelling 
component  OP  that  gives  forward  velocity.  The  line  AB 
is  the  horizontal  ground  line.  If  the  total  lift-drag  ratio 
is  8,  then  the  gliding  angle  will  be  1  in  8,  or  measured  in 
degrees,  tan  0  =  R/W  =  1/8  =  0.125.    From  a  trigonom- 


362 


POWER 


etric  table  it  will  be  found  that  this  tangent  corresponds 
to  an  angle  of  7°  —  10'.  It  should  be  noted  that  R  is  the 
total  resistance  and  not  the  wing-drag. 

Complete  Power  Calculations.  Knowing  the  total 
weight  and  the  desired  speed,  we  must  determine  the  wing 
section  and  area  before  we  start  on  the  actual  power 
calculations.  This  can  either  be  determined  by  empirical 
rules  in  the  case  of  a  preliminary  investigation,  or  by 
actual  calculation  by  means  of  the  lift  coefficients  after 


^ 

\ 

^ 

/  c 

PL>\^fE2 

^ 

^. 

\ 

N^'     "0-GUDING  /^NGLE 

\ 

\ 

^W 

— B 

A 

Fig.  2.  Gliding  Angle  Diagram  Showing  Component  of  Gravity  That  Causes 
Forward  Motion.  The  Gliding  Angle  Depends  Upon  the  Ratio  of 
the  Resistance  to  the  Weight. 


the  approximate  values  are  known.  Sustaining  a  given 
weight,  we  can  vary  the  angle,  area,  wing  section,  or  the 
speed,  the  choice  of  these  items  being  regulated  principally 
by  the  power.  Given  a  small  area  and  a  great  angle  of 
incidence,  we  can  support  the  load,  but  the  power  con- 
sumption will  be  excessive  because  of  the  low  value  of 
the  L/D  ratio  at  high  angles.  If  small  area  is  desired,  a 
large  value  of  Ky  due  to  a  high  lift-wing  section  is  pref- 
erable to  a  low  lift  wing  at  high  angles.  In  general,  the 
area  should  be  so  arranged  that  the  wing  is  at  the  angle  of 
the  maximum  lift-drag  ratio  at  the  rated  speed.    A  low 


POWER  363 

angle  means  a  smaller  motor,  less  fuel,  and  hence  a  lighter 
machine.  This  selection  involves  considerable  difficulty, 
and  a  number  of  wing  sections  and  areas  must  be  tried  by 
the  trial  and  error  method  until  the  most  economical  com- 
bination is  discovered. 

The  first  consideration  being  the  total  weight,  we  must 
first  estimate  this  from  the  required  live  load.  This  can 
be  estimated  from  previous  examples  of  nearly  the  same 
type.  Say  that  our  required  live  load  is  660  pounds,  and 
that  a  live  load  factor  of  0.30  is  used.  The  total  weight 
now  becomes  660/0.30  =  2200  pounds.  To  make  a  pre- 
liminary estimate  of  the  area  we  must  find  the  load  per 
square  foot.  An  empirical  formula  for  biplane  loading 
reads  :  w  =  0.065  V  —  0.25  where  V  =  maximum  speed  in 
M.  P.  H.,  and  w  =  load  per  square  foot.  If  we  assume  a 
maximum  speed  of  90  M.  P.  H.  for  our  machine,  the  unit 
loading  is  w  =  (0.065  x  90)  —  0.25  =  5.6  pounds  per 
square  foot.  The  approximate  area  can  now  be  found 
from  2200/5.6  =  393  square  feet.  (Call  390.)  The  mini- 
mum speed  is  about  48  per  cent  of  the  maximum,  or  43  M. 
P.  H.  We  can  now  choose  one  or  more  wing  sections  that 
will  come  approximately  to  our  requirements  by  the  use  of 
the  basic  formula,  Ky  =  w/V". 

At  high  speed,  Ky  =  5.6/(90  x  90)  =  0.000691.  At  low 
speed,  Ky  =  5.6/(40x40)  =0.003030.  We  must  choose 
the  most  economical  wing  between  these  limits  of  lift,  and 
on  reference  to  our  wing  section  tables  we  find  : 

Wing  Section               Low  Speed  Values.                   High  Speed  Values. 

T>  A  t:.   ^^  umber                 (i)            L/D               Ky                (i)  L/D  Ky 

^•A-F.-6    14°            9.28         0.003018  0°-30'  10.5  0.000690 

RA.F.-3    10°          12.00         0.003000  l°-30'           6.5  0.000690 

E>ffel-32    16°            4.10         0.002908  0°-30'           9.2  0.000690 

U.S.A.-l    14°         10.40         0.003165  0°-00'  11.0  0.000721 

It  would  seem  from  the  above  that  the  chosen  area  is  a 
little  too  large,  as  the  majority  of  the  L/D  ratios  at  high 
speed  are  poor,  the  best  being  11.00  of  the  U.  S.  A.-l.  The 
angles  are  small,  being  negative  in  most  cases  at  high 
speed.    While  the  lift-drag  of  the  R.  A.  F.-3  is  very  good 


36-i  POWER 

at  low  speed,  it  is  very  poor  at  high,  hence  the  area  for  this 
section  should  be  reduced  to  increase  the  loading.  The 
R.  A.  F.-6  and  the  U.  S.  A.-l  show  up  the  best,  for  they 
are  both  near  the  maximum  lift  at  low  speed  and  have  fair 
L/D  ratios  at  high  speed.  It  will  be  seen  that  for  the  best 
results  there  should  be  a  series  of  power  curves  drawn  for 
the  various  wings  and  areas.  This  method  is  too  compli- 
cated and  tedious  to  take  up  here,  and  so  we  will  use  U. 
S.  A.-l,  which  does  not  really  show  up  so  bad  at  this  stage. 
Both  the  R.  A.  F.-6  and  the  U.  S.  A.-l  have  been  used 
extensively  on  machines  of  the  size  and  type  under  con- 
sideration. While  we  require  Ky  =  0.003030,  and  U.  S. 
A.-l  gives  0.003165,  we  will  not  attempt  to  utilize  this 
excess,  as  it  will  be  remembered  that  we  should  not 
assume  the  maximum  lift  for  reasons  of  stability. 

The  wing-drag  at  high  speed  will  be  2200/11.0  =  200 
pounds,  and  at  low  speed  it  will  be:  2200/10.4  =  211 
pounds.  Since  the  maximum  L/D  is  17.8  at  3°,  where  Ky 
is  0.00133,  the  least  drag  will  be  :  2200/17.8  =  124  pounds. 
This  least  drag  will  occur  at  V  =  V5.6/0.00133  =  65  M. 
P.  H. 

The  wing  drag  for  each  speed  must  now  be  divided  by 
the  correction  factor  0.85,  which  converts  the  monoplane 
values  of  drag  into  biplane  values.  Since  this  is  prac- 
tically constant  it  does  not  affect  the  relative  values  of 
Kx  in  comparing  wings,  but  it  should  be  used  in  final 
results. 

For  this  type  of  machine  we  will  take  the  total  parasitic 
resistance  as  r  =  0.036V-.  At  90  M.  P.  H.,  r  =  0.036  x  90 
X  90  =  291.6  pounds.  At  65  M.  P.  H.,  the  resistance  is: 
0.036  X  65  x  65  =  152.1.  At  the  extreme  low  speed  of  43  M. 
P.H.  we  have  r =0.036  x  43  x  43  =  66.56  pounds.  The  total 
resistance  (R)  is  equal  to  the  sum  of  the  wing-drag  and  the 
parasitic  resistance.  At  90  M.  P.  H.  the  total  resistance 
becomes  200  +  291.6  =  491.6  pounds.  At  65  M.  P.  H.  the 
total  is  124  +  152.1  =  176.1,  and  at  43  M.  P.  H.  it  is  211  -f 


I   i  fi   %    M  ^ 


1 

1 

1 

1 

y 

i 

1       i 

1    \  1 

V,    ° 

X 

IS 

^ij 1 — 

i       It 


POWER  365 

66.56^  277.56  pounds.  The  horsepower  is  computed  from 
H  =  RV/375e,  and  at  90  M.  P.  H.  this  is  :  H  =  491.6  x  90/ 
375 x0.80  =  147.5  H.  P.  where  0.80  is  the  assumed  propeller 
efficiency.  At  65  M.  P.  H.  the  horsepower  drops  to  H  ^ 
176.1  X  65/375  x  0.8  =  38.1  H.  P.,  assuming  the  same 
efficiency.  In  the  same  way  the  H.  P.  at  43  M.  P.  H. 
is  39.8. 

A  table  and  power  chart  should  be  worked  out  for  a 
number  of  sections  and  areas  according  to  the  following 
table.  The  calculations  should  be  computed  at  intervals 
of  5  M.  P.  H.,  at  least  the  lower  speeds.  Wing  drag  is  not 
corrected  for  biplane  interference : 


Speed 
M.P.H. 

Lift 

Angle 

L/D 

Drag 

Para.  Res. 

Total  Res.      Total 

(Kv) 

(i) 

(D) 

(R) 

(R)      H.P.(H) 

43 

0.003031 

14" 

10.4 

211 

66.56 

m.'bd             39.8 

65 

0.001333 

3° 

17.8 

124 

152.10 

176.10             38.1 

90 

0.000721 

0° 

11.0 

200 

291.60 

491.60           147.5 

Weight  and  Power.  The  weight  lifted  per  horsepower 
varies  in  the  different  types  of  aeroplanes,  this  difference 
lying  principally  in  the  reserve  allowed  for  climbing  and 
horizontal  speed.  A  speed  scout  may  carry  as  little  as  8 
pounds  per  horsepower,  while  a  slow  two-seater  may 
exceed  20  pounds  per  horsepower.  A  rough  estimate  of 
the  horsepower  required  may  be  had  by  dividing  the  total 
weight  by  the  weight  per  horsepower  ratio  for  that  par- 
ticular type.  Thus  if  the  unit  H.  P.  loading  is  16  pounds 
and  the  total  weight  is  3200,  then  the  horsepower  will 
equal  3200/16  =  200  horsepower.  Assuming  that  the  live 
load  w'  is  0.32  of  the  total  weight  W,  then  W  =  w70.32. 
If  m  =  lbs.  per  H.  P.,  then  H  =  W/m  or  H  =  w70.32m. 
Taking  the  case  of  a  training  machine  where  m  =  20,  and 
the  live  load  is  640  pounds,  the  approximate  horsepower 
will  be :  H  =  w70.32m  =  640/0.32  x  20  =  100  horse- 
power. A  speed  scout  carrying  320  pounds  useful  load, 
with  m  =  10,  will  require  H  =  320/0.32  x  10  =  100  horse- 
power. 


CHAPTER  XVIII. 
PROPELLERS. 

Principles  and  Use  of  Propellers.  A  propeller  converts 
the  energy  of  the  engine  into  the  thrust  required  to  over- 
come the  resistance  of  the  aeroplane.  To  maintain  flight 
the  thrust,  or  force  exerted  by  the  propeller,  must  always 
equal  the  total  resistance  of  the  aeroplane.  A  total  resist- 
ance of  400  pounds  requires  a  propeller  thrust  of  400,  and 
as  the  resistance  varies  with  the  speed,  the  engine  revolu- 
tions must  be  altered  correspondingly.  The  propeller  is 
the  most  complicated  and  least  understood  element  of  the 
aeroplane,  and  we  can  but  touch  only  on  the  most  ele- 
mentary features.  The  inclined  blades  of  the  propeller 
throw  back  an  airstream,  the  reaction  of  which  produces 
the  thrust.  The  blades  can  also  be  considered  as  aerofoils 
moving  in  a  circular  path,  the  lift  of  the  aerofoils  corre- 
sponding to  the  thrust  of  the  propeller.  The  reactions  in 
any  case  are  quite  complicated  and  require  the  use  of 
higher  mathematics  for  a  full  understanding. 

Pitch  and  Velocity.  When  in  action  the  propeller  ro- 
tates, and  at  the  same  time  advances  along  a  straight  line 
parallel  to  its  axis.  As  a  result,  the  tips  of  the  propeller 
blades  describe  a  curve  known  as  *'Helix"  or  screw-thread 
curve.  The  action  is  very  similar  to  that  of  a  screw  being 
turned  in  a  nut.  For  clearness  in  explanation  we  will  call 
the  velocity  in  the  aeroplane  path  the  ''Translational  ve- 
locity," and  the  speed  of  the  tips  in  their  circular  path  as 
the  ''Rotational  velocity."  When  a  screw  works  in  a  rigid 
nut  it  advances  a  distance  equal  to  the  *Titch"  in  each 
revolution,  the  pitch  of  a  single  threaded  screw  being 

366 


PROPELLERS  367 

equal  to  the  distance  between  the  threads.  Since  the 
propeller  or  "Air  screw"  works  in  a  fluid,  there  is  some 
slip  and  the  actual  advance  does  not  correspond  to  the 
*Titch"  of  the  propeller  blades.  The  effective  pitch  is  the 
distance  traveled  by  the  propeller  in  one  revolution.  The 
actual  pitch  or  the  angle  of  the  blades  must  be  greater  than 
the  angle  of  the  effective  helix  by  the  amount  of  slip. 

If  N  =  Revolutions  per  minute,  P  =  eft'ective  pitch  in 
feet  and  V  =  translational  velocity  in  miles  per  hour,  then 
V=NP/88.  With  an  effective  pitch  of  5  feet,  and  1200 
revolutions  per  minute,  the  translational  velocity  of  the 
aeroplane  will  be  :  V  =  1200  x  5/88  =  68.2  miles  per  hour. 


Excelsior    Propeller,    an    Example    of    American    Propeller    Construction.     This 
Propeller  Is  Built  Up  of  Laminations  of  Ash. 

The  actual  pitch  of  the  blades  would  be  from  15  to  25  per 
cent  greater  than  the  effective  pitch  because  of  the  slip.  To 
have  thrust  we  must  have  slip.  With  the  translational 
velocity  equal  to  the  blade-pitch  velocity,  there  is  no  air- 
stream  accelerated  by  the  blades,  and  consequently  there 
is  no  thrust  due  to  reaction.  The  air  thrown  to  the  rear  of 
a  propeller  moves  at  a  greater  speed  than  the  translation 
when  thrust  is  developed,  and  this  stream  is  known  as  the 
''Slipstream."  The  difference  between  the  translational 
and  slipstream  velocity  is  the  slip. 

The  angle  of  the  blade  face  determines  the  pitch.  The 
greater  the  angle  of  the  blade  with  the  plane  of  propeller 
rotation,  the  greater  is  the  pitch.  This  angle  is  measured 
from  the  chord  of  the  working  face  of  the  table,  or  from 
that  side  faced  to  the  rear  of  the  blade.  In  the  majority 
of  cases  the  working  face  is  flat.  The  front  face  is  always 
heavily  cambered  like  a  wing  section,  with  the  greatest 


368  PROPELLERS 

thickness  about  one-third  the  chord  from  the  entering 
edge.  As  in  the  case  of  the  wing,  the  camber  is  of  the 
greatest  importance. 

A  uniform  pitch  propeller  has  a  varying  blade  angle, 
smallest  at  the  tip  and  increasing  toward  the  hub.  With  a 
uniform  pitch  propeller,  every  part  of  the  blade  travels 
through  the  same  forward  distance  in  one  revolution,  hence 
it  is  necessary  to  increase  the  angle  toward  the  hub  as  the 
innermost  portions  travel  a  smaller  distance  around  the 
circle  of  rotation.  Theoretically,  the  angle  at  the  exact 
center  would  be  90  degrees.  The  blade  angles  at  the  dif- 
ferent points  in  the  length  of  a  uniform  pitch  propeller 
are  obtained  as  follows :  Draw  a  right  angle  triangle  in 
which  the  altitude  is  made  equal  to  the  pitch,  and  the  base 
is  equal  to  3.1416  times  the  propeller  diameter.  The  angle 
made  by  the  hypotenuse  with  the  base  is  the  blade  angle 
at  the  tip.  Divide  the  base  into  any  number  of  equal 
spaces  and  connect  the  division  points  with  the  upper 
angle.  The  angles  made  by  these  lines  with  the  base  are 
the  angles  of  the  different  blade  sections. 

Blade  Form.  The  blade  may  be  either  straight-sided  or 
curved.  In  the  latter  case  the  most  deeply  curved  edge  is 
generally  the  entering  edge,  and  the  maximum  width  is 
about  one-third  from  the  tip.  Much  care  is  exercised  in 
arranging  the  outHne  so  that  the  center  of  pressure  will 
not  be  located  in  an  eccentric  position  and  thus  harmfully 
distort  the  blade  when  loaded.  If  this  is  not  attended  to, 
the  pitch  will  vary  according  to  the  load.  In  one  make  of 
propeller  the  blade  is  purposely  made  flexible  so  that  the 
pitch  will  accommodate  itself  correctly  to  different  flight 
speeds  and  conditions.  This,  however,  is  carefully  laid  out 
so  that  the  flexure  is  proportional  throughout  the  blade  to 
the  changes  in  the  load. 

Propeller  Diameter.  The  largest  propellers  are  the 
most  efiicient.  The  propeller  should  be  as  large  as  can  be 
safely  swung  on  the  aeroplane.     Large,  slow  revolution 


PROPELLERS 


369 


^:^ 


O  u 


H13 

a. 


r- 


u 
'^1 


IT,    ^ 


III 


I 


f^JS 


HZ 


3^ 


5  o 
>c/3 


370  PROPELLERS 

propellers  are  far  superior  to  the  small  high  speed  type. 
It  is  more  economical  to  accelerate  a  large  mass  of  air 
slowly  with  a  large  diameter  than  to  speed  up  a  small 
mass  to  a  high  velocity.  The  diameter  used  on  any  aero- 
plane depends  upon  the  power  plant,  propeller  clearance, 
height  of  chassis  and  many  other  considerations.  Ap- 
proximately the  diameter  varies  from  about  1/3  the  span 
on  small  speed  scouts,  to  1/5  or  1/6  of  the  span  on  the 
larger  machines. 

Air  Flow.  The  greater  part  of  the  air  is  taken  in 
through  the  tips,  and  is  then  expelled  to  the  rear.  This 
condition  prevails  until  the  blade  angle  is  above  45  de- 
grees, and  from  this  point  the  flow  is  outward.  Owing  to 
the  great  angles  at  the  hub,  there  is  little  thrust  given 
by  the  inner  third  of  the  blade,  the  air  in  this  region  being 
simply  churned  up  in  a  directionless  mass  of  eddies.  At 
the  tips  the  angle  is  small  and  the  velocity  high,  which 
results  in  about  80  per  cent  of  the  useful  work  being  per- 
formed by  the  outer  third  of  the  blade.  In  some  aero- 
planes a  spinner  cap  is  placed  around  the  hub  to  reduce 
the  churning  loss  and  to  streamline  the  hub.  The  blade 
section  is  very  thick  at  the  hub  for  structural  reasons. 

The  "Disc  area"  of  a  propeller  is  the  area  of  the  circle 
swept  out  by  the  blades.  It  is  the  pressure  over  this  area 
that  gives  the  thrust,  and  in  some  methods  of  calculation 
the  thrust  is  based  on  the  mean  pressure  per  square  foot 
of  disc  area.  The  pressure  is  not  uniformly  distributed 
over  the  disc,  being  many  times  greater  at  the  outer  cir- 
cumference than  at  the  hub.  The  average  pressure  per 
square  foot  depends  upon  the  blade  section  and  angle. 
Because  of  the  great  intensity  of  pressure  at  the  circum- 
ference, the  effective  stream  is  in  the  form  of  a  hollow 
tube. 

Number  of  Blades.  For  training,  and  ordinary  work, 
two-bladed  propellers  are  preferable,  but  for  large  motors 
where  the  swing  is  limited,  three  or  four  blades  are  often 


PROPELLERS  371 

used.     A  multlple-bladed  propeller  absorbs  more  horse- 
power with  a  given  diameter  than  the  two-blade  type.    In 
general,  a  four-bladed  propeller  revolving  slowly  may  be 
considered  more  efficient  than  the  two-blade  revolving  # 
rapidly.    Where  the  swing  and  clearance  are  small,  a  small  / 
four-blade  may  give  better  results  than  a  larger  and  faster!    9 
two-blade.    A  three-blade  often  shows  marked  superiority  ! 
over  a  tw^o-blade  even  when  of  smaller  diameter,  and  the  ; 
hub  of  the  three-blade  is  much  stronger  than  the  four- 1 
blade,  although  neither  the  three  or  four  is  as  strong  as  I 
the  two-blade  type. 

Effects  of  Altitude.  At  high  altitudes  the  density  is 
less,  and  consequently  the  thrust  is  less  with  a  given  num- 
ber of  revolutions  per  minute.  The  thrust  can  be  main- 
tained either  by  increasing  the  speed,  or  by  increasing  the 
pitch.  For  correct  service  at  high  altitudes  the  propeller 
should  undoubtedly  be  of  the  variable  pitch  type,  in  which 
the  pitch  can  be  controlled  manually,  or  by  some  auto- 
matic means  such  as  proportional  blade  flexure. 

Effects  of  Pitch.  Driven  at  a  constant  speed,  both  the 
thrust  and  horsepower  increase  with  the  pitch  up  to  a 
certain  limiting  angle. 

For  a  given  horsepower  the  static  thrust  depends  both 
on  the  diameter  and  the  pitch.  If  the  pitch  is  increased 
the  diameter  must  be  decreased  in  proportion  to  maintain 
a  constant  speed.  As  the  pitch  is  regulated  by  the  trans- 
lational  speed  and  revolutions,  the  static  thrust  of  a  high 
speed  machine  is  very  small.  As  the  translational  speed 
increases,  the  pitch  relative  to  the  wind  is  less,  and  conse- 
quently the  thrust  will  pick  up  until  a  certain  limiting 
speed  is  reached. 

Thrust  and  Horsepower.  The  calculation  for  thrust 
and  power  are  very  complicated,  but  the  primary  condi- 
tions can  be  given  by  the  following:  Let  V  =  the  pitch 
velocity  in  feet  per  minute,  T  =  thrust  in  pounds,  and  H 
=  horsepower,  then  H  =  TV/33000E  from  which  T  = 


372  PROPLLLERS 

33000HE/V,  the  efficiency  being  designated  by  E.  Since 
the  pitch  velocity  is  NP,  where  N  =  revs,  per  minute  and 
P  =  pitch  in  feet,  then  T  =  33000HE/PN.  Assuming  a 
5-foot  pitch,  1200  revs.,  the  efficiency  =  0.75,  and  the 
horsepower  100,  the  thrust  will  be : 

T  =  33000  X  100  X  0.75/5  x  1200  =  412.5  pounds.  The 
pitch  in  this  case  is  the  blade  pitch,  and  the  great  uncer- 
tainty lies  in  selecting  a  proper  value  for  E.  This  may 
vary  from  0.70  to  0.85.  The  diameter  is  also  an  unknown 
factor  in  this  primitive  equation. 

Materials  and  Construction.  The  woods  used  for  pro- 
peller construction  are  spruce,  ash,  mahogany,  birch, 
white  oak,  walnut,  and  maple.  Up  to  50  H.  P.  spruce  is 
suitable,  as  it  is  light,  and  strong  enough  for  this  power. 
In  Europe  walnut  and  mahogany  are  the  most  commonly 
used,  although  they  are  very  expensive.  Birch  is  very 
strong  and  comparatively  light  for  its  strength,  and  can 
be  used  successfully  up  to  125  horsepower.  Ash  is  strong, 
light  and  fibrous,  but  has  the  objectionable  feature  of 
warping  and  cannot  withstand  moisture.  Maple  is  too 
beavy  for  its  strength.  White  oak,  quarter-sawed,  is  the 
best  of  propeller  woods  and  is  used  with  the  very  largest 
engines.  It  is  strong  for  its  weight  and  is  hard,  but  is 
very  difficult  to  work  and  glue.  For  tropical  climates, 
Southern  poplar  is  frequently  used  as  it  has  the  property 
of  resisting  heat  and  humidity. 

One-inch  boards  are  rough  dressed  to  %  inch  and  then 
finished  down  to  ^f  or  ^  inch.  After  a  thorough  tooth 
planing  to  roughen  the  surface  for  the  glue,  they  are  thor- 
oughly coated  with  hot  hide  glue,  piled  together  in  blocks 
of  from  5  to  10  laminations,  and  then  thoroughly  squeezed 
for  18  hours  in  a  press  or  by  clamps  until  the  glue  has 
thoroughly  set.  Only  the  best  of  hide  glue  is  used,  applied 
at  a  temperature  of  140°  F.  and  at  a  room  temperature  of 
100°.  The  glue  must  never  be  hotter,  nor  the  boards 
cooler  than  the  temperatures  stated.    The  propeller  after 


PROPELLERS 


373 


being  roughed  out  is  left  to  dry  for  ten  days  so  that  all 
of  the  glue  stresses  are  adjusted.  If  less  time  is  taken, 
the  propeller  will  warp  out  of  shape.  The  propeller  is 
worked  down  within  a  small  fraction  of  the  finished  size 
and  is  again  allowed  to  rest.  After  a  few  days  it  is 
finished  down  to  size  by  hand,  is  scraped,  and  tested  for 
pitch,  tracking  and  hub  dimensions. 

The  finish  is  glossy,  and  may  be  accomplished  by  sev- 
eral coats  of  spar  varnish  or  by  repeated  applications  of 
hot  boiled  linseed  oil  well  rubbed  in,  finishing  with  three 
or  four  coats  of  wax  polish.  There  should  be  at  least  5 
applications  of  linseed  oil,  the  third  coat  being  sandpapered 
with  No.  0  paper.  The  wood  should  be  scraped  to  dimen- 
sion and  must  not  be  touched  with  sandpaper  until  at  least 
two  coats  of  varnish  or  oil  have  been  applied. 

The  wood  must  be  absolutely  clear  and  straight  grained, 
and  without  discolorations.  The  boards  must  be  piled 
so  that  the  edge  of  the  grain  is  on  the  face  of  the  blades, 
and  the  direction  of  the  annular  rings  must  be  alternated 
in  the  adjacent  boards. 


Houleur   ifM  (avec  Mice   3:io) 


374 


PROPELLERS 


CHAPTER  XIX. 
OPERATION  AND  TRAINING. 

Self-Training.  In  the  early  days  of  aviation,  there  were 
few  schools,  and  these  were  so  expensive  to  attend  that 
the  majority  of  the  aeronautical  enthusiasts  taught  them- 
selves to  fly  on  home-made  machines.  While  this  was  a 
heroic  method,  it  had  the  advantage  of  giving  the  student 
perfect  confidence  in  himself,  and  if  his  funds  were  suffi- 
cient to  outlast  the  crashes,  it  resulted  in  a  finished  and 
thorough  flyer.  In  general,  this  process  may  be  described 
as  consisting  of  two  hours  of  practice  followed  by  a  week 
or  more  of  repairing. 

The  present-day  beginner  has  many  advantages.  He 
has  the  choice  of  many  excellent  schools  that  charge  a 
reasonable  tuition,  and  where  the  risk  of  injury  is  small. 
He  has  access  to  the  valuable  notes  published  in  the  aero- 
nautical magazines,  and  the  privilege  of  consulting  with 
experienced  aviators.  The  stability  and  reliability  of  the 
planes  and  the  motors  have  also  been  improved  to  a  re- 
markable degree,  and  the  student  no  longer  has  to  con- 
tend with  a  doubtful  aeroplane  construction  nor  with  the 
whims  of  a  poorly-constructed  motor. 

Training  Methods.  In  the  majority  of  American 
schools,  the  instructor  accompanies  the  student  in  the 
first  flights.  The  controls  are  *'Dual,"  or  interconnected, 
so  that  the  instructors'  controls  act  in  unison  with  those 
of  the  student,  thus  giving  the  latter  an  accurate  knowl- 
edge of  the  movements  necessary  for  each  flight  condition. 
After  the  first  few^  flights  the  instructor  can  relax  his  con- 
trols at  times  so  that  the  student  can  take  charge.  This 
continues  until  the  student  has  shown  the  ability  to  handle 

375 


376  TRAINING  AND  OPERATION 

the  machine  alone  under  ordinary  conditions  and  is  then 
ready  for  his  first  "Solo"  or  flight  alone.  The  first  solo  is 
a  critical  period  in  his  training,  for  when  once  in  flight  he 
is  beyond  all  human  aid. 

At  the  navy  training  school  at  Pensacola,  the  student  is 
first  taken  for  a  ride  with  one  of  the  instructors  without 
giving  him  access  to  the  controls.  This  is  simply  to 
give  the  student  an  experience  in  the  sensation  of  flight. 
After  this  he  is  taken  for  a  series  of  short  flights  on  a 
dual  control  machine,  the  instructor  gradually  allowing 
him  to  take  charge  to  a  greater  and  greater  extent  as  he 
develops  the  "Air  feel."  During  this  time  the  intricacies 
of  the  maneuvers  are  also  gradually  increased,  so  that 
after  about  ten  hours  of  this  sort  of  work  he  is  allowed 
to  take  his  first  solo.  It  has  been  found  that  the  average 
student  will  require  from  10  to  20  hours  of  dual  control 
instruction  before  he  is  fit  to  fly  alone.  When  his  work 
has  proven  satisfactory  he  is  then  allowed  to  fly  in  rough 
weather,  execute  spirals,  and  attempt  high  altitude  and 
long  distance  flying. 

Some  instructors  believe  in  showing  what  can  be  per- 
formed in  the  air  from  the  very  beginning.  During  the 
first  dual  flights,  the  pilot  indulges  in  dives,  vertical  banks, 
side  slip,  or  even  looping.  After  an  experience  of  this 
sort,  the  student  is  far  more  collected  and  easy  during 
O  the  following  instructions  in  simple  straight  flying.  If 
f  this  preliminary  stunt  flying  has  a  very  material  effect  on 
the  nerve  of  the  student  it  may  be  taken  for  granted  that 
he  is  not  adapted  for  the  work  and  can  be  weeded  out 
without  further  loss  of  time.  If  he  is  of  the  right  type, 
this  "rough  stufif"  has  a  beneficial  influence  on  his  work 
during  the  succeeding  lessons.  During  this  time  numer- 
ous landings  are  made,  for  it  must  be  understood  that  this 
is  one  of  the  most  difficult  features  of  flying.  With  15 
minute  lessons,  at  least  6  landings  should  be  made  per 
lesson. 


TRAINING  AND  OPERATION  377 

A  second  method  of  instruction,  and  to  the  author's 
mind  the  most  desirable,  is  by  means  of  the  "Penguin" 
or  "Roller."  This  is  a  low-powered  machine  with  very- 
small  wings — so  small  that  it  cannot  raise  itself  from  the 
ground.  By  running  the  penguin  over  the  ground,  the 
student  learns  how  to  manage  his  engine  and  to  steer  with 
his  feet.  In  this  way  he  obtains  a  certain  delicacy  of 
touch  without  endangering  either  himself  or  an  expensive 
machine.  After  he  has  progressed  satisfactorily  on  this 
machine  he  graduates  to  a  faster  penguin  or  else  to  a 
very  slow  aeroplane  with  which  he  can  actually  leave 
the  ground.  Since  the  second  penguin,  or  the  slow 
aeroplane  are  much  faster  than  the  first  machine,  the 
student  finds  that  the  sensitiveness  of  the  rudder  and 
controls  are  greatly  increased.  They  require  more  careful 
handling  than  in  the  first  instance,  and  the  slightest  mis- 
take or  delay  will  send  the  machine  skidding.  The  aero- 
planes used  at  this  stage  are  very  low-powered,  and  are 
capable  of  rising  only  a  few  feet  from  the  ground,  but  they 
give  the  student  an  opportunity  of  learning  the  aileron  and 
tail  controls  in  comparative  safety.  The  same  result  can 
be  obtained  with  a  standard  aeroplane  by  a  permanent 
set  in  the  throttle  control,  and  by  adjusting  the  stabilizer 
surface.  The  beginner  is  allowed  to  work  only  during 
calm  weather,  as  the  low  speed  and  small  lifting  capacity 
is  likely  to  cause  an  accident  if  the  machine  is  caught  by 
a  side  or  following  gust.  He  only  learns  how  to  get  the 
machine  oflf  the  ground,  to  keep  the  tail  up  and  hold  it 
in  a  straight  line  for  a  few  seconds. 

The  man  taught  by  the  penguin  method  is  alone  when 
he  first  leaves  the  ground,  and  hence  is  generally  more 
self-reliant  than  one  who  has  been  "Spoon  fed."  His  expe- 
rience in  handling  the  controls  has  made  his  movements 
instinctive,  so  that  when  he  first  actually  flies  he  is  in  a 
better  position  to  analyze  the  new  problems  before  him. 
It  is  a  better  and  cheaper  method  for  the  school  as  the 


378  TRAINING  AND  OPERATION 

breakage  is  less  expensive  and  allows  the  unfit  students  to 
be  weeded  out  before  they  cause  damage  to  themselves  or 
to  the  school  property. 

Ground  Instruction.  Before  attempting  flight,  the 
student  should  be  thoroughly  versed  in  the  principles  and 
constructional  details  of  the  aeroplane  and  the  aeronautic 
motor.  He  should  know  how  to  take  down,  time  and 
repair  every  type  of  motor  with  which  he  is  likely  to  come 
into  contact.  He  should  be  able  to  tell  at  a  glance  whether 
the  machine  is  rigged  or  trued  up  properly,  and  have  a 
general  knowledge  of  the  underlying  principles  of  aero- 
dynamics. The  study  of  these  subjects  is  the  function  of- 
the  ground  school.  At  this  school  the  student  should 
learn  the  assembling  and  adjusting  of  the  aeroplane  struc- 
ture and  its  balance. 

Types  Suitable  for  Pilots.  There  is  a  great  diversity  of 
opinion  as  to  the  type  of  man  best  suited  for  flying.  In 
this  country  the  government  requirements  regarding  age 
and  physical  condition  are  very  exacting,  while  in  Europe 
it  has  been  found  that  physical  condition  is  not  an  index  to 
a  man's  ability  as  a  pilot.  Many  of  the  best  French  pilots 
were  in  such  bad  shape  as  to  be  rejected  by  the  other 
branches  of  the  army.  Our  men  are  well  under  30  years  of 
age,  while  in  European  service  there  are  many  excellent 
pilots  well  over  40.  It  is  almost  impossible  to  tell  from 
external  appearances  whether  a  man  can  become  a  good 
pilot. 

In  general  he  must  be  more  intelligent  and  better  edu- 
cated than  the  average  infantryman.  He  should  not  be 
subject  to  an  attack  of  "Nerves,"  nor  become  easily  rat- 
tled, for  such  a  man  courts  disaster  in  flying.  Many  exhi- 
bition flyers  of  reputation  have  proved  absolute  failures 
in  military  service.  A  knowledge  of  mechanics  will  be  of 
great  benefit  and  has  been  the  salvation  of  many  a  pilot 
in  active  service.  Automobile  or  motorcycle  experience  is 
particularly  valuable.     Recklessness,  or  a  dare-devil  sort 


TRAINING  AND  OPERATION  379 

of  a  disposition,  are  farthest  from  being  qualifications  for 
an  aviator.  Such  a  man  should  not  be  permitted  to  fly, 
for  he  is  not  only  a  constant  menace  to  himself  but  to 
everyone  else  concerned. 

Learning  to  Fly  Alone.  It  is  with  the  greatest  hesi- 
tancy that  the  author  enters  into  a  "Ground  course"  of 
flight  instruction.  I  can,  however,  list  the  principal  things 
to  avoid  and  some  of  the  things  to  do,  but  this  will  never 
take  the  place  of  actual  field  instruction  and  experience. 
The  first  and  last  thing  to  remember  is  to  "Proceed  slowly 
and  with  caution."  Never  try  a  new  stunt  until  you  are 
absolutely  sure  that  you  have  thoroughly  mastered  the 
preliminary  steps  in  straight  flying.  Over-confidence  at 
the  beginning  is  almost  as  bad  as  no  confidence  at  all, 
and  the  greatest  difficulty  met  with  by  instructors  during 
the  first  solo  flights  is  to  keep  the  student  from  imitating 
the  maneuvers  of  the  more  experienced  flyers.  Spend 
plenty  of  time  rolling  or  "Grass-cutting"  before  attempt- 
ing to  leave  the  ground.  Be  sure  that  you  can  handle  the 
rudder  with  accuracy,  and  at  fairly  high  speeds  before  at- 
tempting to  Hft.  A  few  days  spent  in  sitting  in  the  ma- 
chine (motor  dead),  and  acquainting  yourself  with  the 
controls  is  excellent  practice  and  certainly  is  not  a  loss 
of  time.  With  the  machine  in  the  hangar,  move  the  con- 
trols for  imaginary  turns,  dips  and  other  maneuvers  so 
that  the  resistance,  reach  and  limit  of  control  movement 
will  come  more  naturally  when  the  machine  is  moving. 

During  the  ground  rolling  period,  the  elevator  or  stabil- 
izer should  be  set  so  that  it  is  impossible  to  leave  the 
ground,  and  the  motor  should  be  adjusted  so  that  it  cannot 
develop  its  full  thrust.  This  will  provide  against  an  acci- 
dental lift.  Be  easy  and  gentle  in  handling  the  controls,  for 
they  work  easily,  and  have  powerful  effect  at  high  speeds. 
The  desperate  fervor  with  which  the  beginner  generally 
yanks  at  the  "joy  stick"  is  generally  the  very  reason  for  his 
accidents.     Do  not  start  oflf  at  full  speed  without  first 


380  TRAINING  AND  OPERATION 

getting  use  to  the  effect  of  the  controls.  Learn  to  find 
the  location  of  the  various  devices  so  that  you  can  reach 
them  v^ithout  looking  or  without  fumbHng. 

The  First  Straight.  By  adjusting  the  stabilizer  and  ele- 
vators so  that  the  latter  has  a  greater  degree  of  freedom, 
and  by  changing  the  motor  so  that  it  can  be  run  at  a  slight- 
ly higher  speed,  we  are  in  a  position  to  attempt  our  first 
flight.  Be  careful  that  the  adjustment  will  limit  the  climb 
of  the  machine,  and  choose  only  the  calmest  of  weather.  It 
should  be  remembered  that  the  aeroplane  will  get  off  the 
ground  at  a  lower  speed  than  that  required  for  full  flight 
at  higher  altitudes,  this  being  due  to  the  cushioning 
effect  between  the  wings  and  the  earth.  A  machine  trav- 
eling at  a  speed  capable  of  sustaining  flight  at  a  few  feet 
above  the  earth  will  cause  it  to  stall  when  it  is  high  enough 
to  lose  this  compression.  The  adjustment  should  be  such 
that  the  machine  cannot  rise  above  this  "Cushion,"  and 
in  this  condition  it  is  fairly  safe  for  the  beginner. 

In  making  the  first  runs  under  the  new  conditions  of 
adjustment,  the  student  should  learn  to  manipulate  the 
elevators  so  that  they  will  hold  the  tail  up  in  the  correct 
position,  that  is,  with  the  chord  of  the  wings  nearly  hori- 
zontal. Do  not  allow  the  tail  skid  to  drag  over  the  ground 
further  than  necessary.  At  this  point  the  student  should 
be  strapped  in  the  seat  by  a  quick-detachable  safety 
belt. 

Now  comes  the  test.  Get  under  full  headway  with  the 
tail  well  up,  taking  care  to  run  against  the  breeze.  The 
speed  increases  rapidly,  then  the  motion  and  jar  seem 
softer,  and  the  motor  ceases  to  roar  so  loudly.  There  is 
now  a  very  distinct  change  in  the  note  of  the  motor.  You 
are  off.  At  this  point  a  very  peculiar  illusion  takes  place, 
for  your  elevation  of  a  few  feet  seems  about  a  thousand 
times  greater  than  it  really  is.  With  this  impression  the 
student  usually  tries  to  correct  matter  by  a  sudden  for- 
ward push  on  the  control  lever  causing  fine  dive  and  a 


TRAINING  AND  OPERATION  381 

smash.  It  must  be  borne  in  mind  that  only  the  slightest 
movement  of  the  controls  should  be  made,  and  if  this  does 
not  prove  sufficient  after  a  moment  or  so,  advance  them 
still  further  but  very  gently.  Sudden  movements  must  be 
avoided.  At  first  the  "Hops"  should  not  extend  over  a 
hundred  yards  or  so  until  the  student  is  sure  of  his  con- 
trols. Little  by  little  they  can  be  increased  in  length  and 
height.  He  should  practice  for  some  time  before  attempt- 
ing a  flight  of  more  than  a  mile.  By  this  time,  the  student 
will  have  learned  that  the  landing  is  by  far  the  most  diffi- 
cult feature  in  flying,  and  he  should  practice  this  inces- 
santly before  trying  flights  in  windy  weather. 

The  machine  should  be  headed  directly  into  the  wind, 
both  in  getting  off  and  in  landing,  especially  in  the  latter 
case,  as  a  sudden  following  gust  will  tend  to  stall  a 
machine  or  upset  it.  With  a  head  wind,  the  lift  is  main- 
tained at  a  low  speed  and  hence  is  an  aid  in  a  safe  landing. 
When  flying  in  still  air  there  is  little  if  any  use  for  the 
ailerons,  but  in  gusts  the  student  will  need  their  aid  in 
maintaining  lateral  balance.  After  the  rudder  and  ele- 
vator controls  have  been  well  learned  the  effect  of  the 
ailerons  can  be  tried.  Gusty  or  squally  weather  must  be 
avoided  at  this  point  in  the  training,  and  no  turns  should 
yet  be  attempted. 

When  the  student  attains  heights  greater  than  a  few 
feet  he  should  take  great  care  in  obtaining  a  sufficient 
ground  speed  before  trying  to  get  off,  for  if  lifted  before 
the  full  flying  speed  is  attained  it  is  likely  to  stall.  Fast 
climbing  at  sharp  angles  is  dangerous  unless  a  sufficient 
ground  speed  has  been  attained.  Sustentation  is  due  to 
forward  speed,  and  this  must  not  be  forgotten.  The 
quickest  climb  for  getting  over  trees  and  other  obstruc- 
tions is  obtained  by  gaining  full  speed  on  the  ground  be- 
fore the  climb  begins,  as  the  power  of  the  engine  is  aided 
by  the  momentum  of  the  machine. 

In  landing  in  small  fields  it  is  necessary  to  bring  the 


382  TRAINING  AND  OPERATION 

machine  to  rest  as  soon  as  possible,  and  this  stopping 
distance  depends  to  a  great  extent  upon  the  attitude  of 
the  machine  when  it  first  touches  the  ground.  If  it  is 
landed  so  that  the  chassis  wheels  and  tail  skid  strike  the 
ground  simultaneously,  the  incidence  is  so  great  that  the 
wings  act  as  air  brakes.  On  landing,  the  angle  in  any  case 
should  be  quickly  increased  past  the  angle  of  maximum 
lift.  The  lift  is  much  reduced  and  the  drag  is  increased  by 
quickly  pulling  the  control  toward  the  aviator.  This  also 
reduces  the  tendency  toward  nosing  over. 

A  normal  landing  in  a  large  field  can  be  affected  by  first 
starting  down  at  the  normal  gliding  angle,  and  when  from 
twenty  to  thirty  feet  above  the  ground  the  elevator  con- 
trol is  pulled  back  so  that  the  machine  will  describe  a 
curve  tangent  to  the  ground.  In  student's  practice  the 
curve  should  not  be  exactly  tangent  to  the  ground,  but 
tangent  to  a  level  two  or  three  feet  above  the  ground. 
The  machine  is  now  losing  speed,  and  to  prevent  settling 
the  elevator  should  be  pulled  back  a  trifle.  The  speed 
continues  to  decrease  until  it  settles  down  through  the 
small  remaining  distance  with  the  elevator  full  back.  The 
points  of  support  should  strike  simultaneously.  It  is  dif- 
ficult for  the  beginner  to  make  this  sort  of  a  landing,  as 
there  always  seems  to  be  an  uncontrollable  desire  to  jam 
the  machine  down  on  the  ground.  If  a  puff  of  wind  hap- 
pens to  strike  the  machine  when  a  few  feet  off,  the  student 
becomes  rattled  by  the  suddenly  increased  elevation  and 
jams  her  down  doubly  hard. 

Wind  Flying.  The  nature  of  wind  at  low  altitudes  is 
determined  to  a  great  extent  by  the  contour  of  the  ground. 
Eddies  are  caused  by  trees,  embankments,  fences,  small 
hills,  etc.,  which  tend  to  disturb  the  equilibrium  or  change 
the  course  of  the  aeroplane.  As  the  altitude  increases, 
the  effects  of  these  obstructions  are  less  pronounced,  until 
at  from  2000  to  3000  feet  the  effect  is  practically  negligible. 
Winds  that  may  be  "Bumpy"  near  the  ground  are  fairly 


TRAINING  AND  OPERATION  383 

regular  when  3000  feet  is  attained.  At  the  higher  altitudes 
the  velocity  increases,  and  if  the  machine  is  flying  against 
the  wind  the  progress  will  naturally  be  much  slower  at  the 
higher  altitudes.  When  starting  in  a  strong  wind  it  is 
advisable  to  attain  an  altitude  of  at  least  300  to  400  feet 
before  turning.  Turning  in  with  the  wind  carries  the  pos- 
sibility of  a  drop  or  stall. 

A  short  gust  striking  the  machine,  head  on,  tends  to 
retard  the  velocity  in  regard  to  the  earth,  but  in  reality  in- 
creases the  relative  air  speed  and  thus  causes  the  machine 
to  climb  momentarily.  A  prolonged  head  gust  may  pro- 
duce a  stall  unless  corrected  by  the  elevator  or  met  with 
by  reserve  power.  A  rear  gust  reduces  the  relative  wind 
velocity  and  tends  to  make  the  machine  stall,  although 
there  are  a  few  cases  where  the  gust  velocity  has  been 
great  enough  to  cause  a  precipitate  drop.  The  higher  the 
speed,  the  less  the  danger  from  rear  gusts. 

The  gusts  are  much  more  pronounced  with  low  winds, 
say  winds  of  about  5  to  15  miles  per  hour,  and  hence  it  is 
usually  more  tricky  to  fly  in  a  wind  of  this  velocity  than 
with  a  higher  wind.  It  is  not  the  speed  of  the  wind  so 
much  as  it  is  its  variation  from  the  average  velocity.  One 
should  start  to  work  on  a  ''bump"  at  the  moment  it  first 
starts  to  appear. 

When  flying  with  the  wind,  the  total  speed  in  regard  to 
the  earth  is  the  sum  of  the  wind  speed  and  the  aeroplane 
speed.  When  flying  against  it  is  the  difference  between 
the  aeroplane  and  air  speeds.  Thus,  if  the  air  speed  of  the 
aeroplane  is  60  miles  per  hour,  the  speed  in  regard  to  the 
ground  will  be  7})  miles  per  hour  with  a  following  wind  of 
15  miles  per  hour,  and  45  miles  per  hour  when  flying 
against  a  15-mile  wind.  The  speed  when  flying  across 
the  wind  would  be  represented  by  the  diagonal  of  a  paral- 
lelogram, one  side  of  which  represents  the  aeroplane 
speed,  and  the  other  side  the  wind  speed.  The  angle  of 
the  diagonal  is  the  angle  at  which  the  machine  must  be 


384  TRAINING  AND  OPERATION 

pointed.  When  viewed  from  the  ground,  an  aeroplane  in 
a  cross  wind  appears  to  fly  sideways. 

Turning.  After  the  beginner  is  able  to  maintain  longi- 
tudinal and  lateral  balance  on  straight  away  flights,  he 
next  attempts  turns.  At  first,  the  turns  must  be  of  great 
radius.  As  the  radius  is  gradually  shortened,  the  effects 
of  centrifugal  force  become  greater,  increasing  the 
tendency  toward  skidding  or  outward  side  slip.  To  pre- 
vent skidding,  the  outer  wing  tip  must  be  raised  so  that 
the  lift  will  oppose  the  centrifugal  force.  The  shorter  the 
turn,  and  the  faster  it  is  made,  the  greater  will  be  the 
banking  angle.  Should  the  bank  be  too  steep,  the  gravita- 
tional force  will  pull  the  machine  down,  and  inwardly  in 
a  direction  parallel  to  the  wings.  This  is  known  as  an 
*Tnner  side  slip."  The  banking  may  be  performed  by  the 
natural  banking  tendency  of  the  aeroplane  or  may  be 
assisted  by  depressing  the  aileron  on  the  outer  wing  tip. 
Unless  the  speed  is  well  up  to  normal,  the  machine  will  be 
likely  to  stall  and  drop  on  a  turn,  as  the  head  resistance 
is  much  greater  under  these  conditions.  For  safety  one 
should  take  a  short  downward  gHde  before  starting  the 
turn,  so  that  the  speed  will  surely  be  sufficient  to  carry  it 
around  the  turn.  A  turn  should  never  be  attempted  when 
climbing  unless  one  has  a  great  reserve  power.  The  com- 
bined effects  of  the  turning  resistance,  and  absorption  of 
energy  due  to  the  climb,  will  be  almost  certain  to  stall  the 
machine.  There  are  banking  indicators  on  the  market 
which  will  prove  of  great  service.  These  operate  on  the 
pendulum  principle  and  indicate  graphically  whether  the 
aeroplane  is  being  held  at  the  correct  angle  of  bank. 

Proper  Flight  Speed.  An  aeroplane  should  always  be 
provided  with  an  air  speed  meter,  giving  the  speed  of  the 
machine  in  relation  to  the  air.  When  flying  with  the  wind 
the  pilot  is  likely  to  be  confused  by  the  tremendous  ground 
speed  at  which  his  machine  is  flying.  While  the  machine 
may  be  moved  at  a  fast  clip  in  regard  to  the  earth,  it  may 


TRAINING  AND  OPERATION  385 

be  really,  near  the  stalling  speed.  The  pilot's  sense  of 
speed  is  influenced  by  the  rapidly  moving  objects  below 
him. 

Spinning  Nose  Dive.  Most  pilots  sooner  or  later  get 
into  a  spinning  nose  dive,  either  throtgh  side  slip  or  an 
incorrectly  designed  machine.  If  the  machine  is  spirally 
unstable,  it  is  almost  certain  to  get  into  a  spinning  nose 
dive  when  it  is  stalled  on  a  sharp  turn  or  is  sharply  turned 
at  a  low  speed.  What  first  appears  to  be  a  side  slip  then 
starts.  The  nose  will  drop,  and  the  tail  will  start  to  swing 
around  in  a  larger  circle  than  the  nose.  Unless  immediate- 
ly corrected  this  will  end  in  a  bad  smash.  If  the  spin  is 
opposite  to  the  rotation  of  the  motor,  the  motor  should  be 
stopped  since  the  torque  helps  in  maintaining  the  spin. 

Push  the  elevators  into  a  diving  position  and  turn  the 
rudder  against  the  direction  of  the  spin.  If  this  does  not 
break  up  the  rotation,  wait  until  the  elevators  have 
brought  the  machine  into  a  vertical  diving  position  and 
then  put  the  ailerons  hard  over  against  the  spin  direction. 
When  the  ailerons  have  stopped  the  spin,  the  dive  can  be 
straightened  out  by  the  elevators  in  the  ordinary  way. 
The  controls  may  not  take  hold  immediately,  but  this  is 
no  reason  for  changing  the  method  of  procedure. 

Straight  Nose  Dive.  When  the  aeroplane  is  diving 
vertically,  nose  down,  the  center  of  pressure  movement 
sometimes  opposes  the  elevators  and  makes  it  difficult  to 
straighten  out  of  the  dive.  Naturally,  the  aviator  will 
have  the  elevator  tip  pointed  up,  as  in  the  case  of 
straightening  out  from  a  glide,  but  if  this  is  not  effective 
in  changing  the  attitude  he  should  throw  the  elevator  to 
the  extreme  opposite  position  for  a  moment.  This  will 
tend  to  swing  the  machine  over  on  its  back  and  relieve  the 
pressure  on  the  main  surface.  A  quick  movement  of  the 
elevator  to  its  original  position  will  give  control. 

Gliding  or  Volplane.     Gravity  may  be  made  to  give  a 
propelling  component  by  pointing  the  nose  down  along  an 


386  TRAINING  AND  OPERATION 

inclined  "Gliding  path."  In  many  ways  this  can  be  com- 
pared to  a  weight  sliding  down  an  inclined  plane.  The 
smaller  the  angle  made  with  the  horizontal,  the  greater 
will  be  the  radius  of  action,  and  the  better  will  be  the 
opportunity  of  picking  out  a  safe  landing  place.  This 
angle  is  an  inherent  feature  of  every  aeroplane  and  bears 
a  direct  relation  to  the  weight  and  the  head  resistance. 
Very  efficient  machines  will  have  a  glide  of  about  1  in  12, 
or  will  travel  12  feet  horizontally  while  dropping  1  foot. 
At  an  elevation  of  2000  feet  this  aeroplane  will  travel  for- 
ward a  distance  equal  to  2000  x  12  =  24000  feet  before 
touching  the  ground. 

Imimebnan  Turn.  This  is  a  "Stunt"  devised  by  the  Ger- 
man flyer  Immelmann  for  evading  enemy  machines.  It 
allows  the  aeroplane  to  be  driven  straight  at  the  enemy 
machine  and  then  directly  away  from  it.  The  machine  is 
first  rolled  over  on  its  back  in  a  sideways  rotation,  and  is 
then  brought  into  a  dive.  When  straightened  out,  right 
side  up,  the  machine  will  be  scooting  away  at  full  speed 
in  the  opposite  direction.  This  is  a  very  quick  and 
effective  method  of  dodging  fire  from  the  enemy  guns. 


CHAPTER  XX. 
AERONAUTICAL  MOTORS. 

General  Notes.  It  is  assumed  that  the  reader  under- 
stands the  principles  of  the  automobile  motor  and  its  ac- 
cessories, for  a  minute  description  of  gas-engine  principles 
does  not  fall  within  the  scope  of  this  book.  If  more  in- 
formation is  desired  on  this  subject,  the  reader  is  referred 
to  the  author's  "Practical  Handbook  of  Gas,  Oil  and  Steam 
Engines."  Only  those  features  pecuHar  to  aeronautic 
motors  will  be  discussed  in  this  chapter. 

Aeronautic  Requirements.  The  principal  requirements 
of  an  aeronautic  motor  are  light  weight,  low  oil  and  fuel 
consumption,  reliability  and  compactness.  The  outline  as 
viewed  from  the  shaft  end  is  also  very  important,  for  the 
motor  must  be  mounted  in  a  narrow  streamline  body.  The 
compression  pressures  are  much  higher  than  those  em- 
ployed on  auto  motors,  and  the  speed  is  generally  lower. 
With  one  or  two  exceptions  the  four-stroke  cycle  has  been 
universally  adopted. 

Aeronautic  service  is  a  severe  test  for  the  motor. 
From  the  start  to  the  finish  of  a  flight,  the  aero- 
plane motor  is  on  a  steady  grind,  loaded  at  least  to  75  per 
cent  of  its  rated  power.  The  foundations  are  light  and 
yielding  and  the  air  density  varies  rapidly  with  changes  in 
the  altitude.  As  the  fuel  and  oil  require  an  expenditure 
of  power  for  their  support,  the  fuel  consumption  becomes 
of  great  importance,  especially  in  long  flights.  Because 
of  the  heavy  normal  load  the  lubricating  system  must  be 
as  nearly  perfect  as  it  is  possible  to  make  it. 

A  motor  car  runs  normally  at  from  10  to  25  per  cent 
of  its  rated  horsepower,  while  the  aero  motor  may  develop 

387 


388  MOTORS 

as  high  as  75  per  cent  to  100  per  cent  for  hours  at  a  time. 
A  car  engine  of  672  cubic  inches  displacement  is  rated  at 
65  horsepower,  while  the  same  size  aero  engine  has  a 
rating  of  154.  On  the  basis  of  normal  output,  this  ratio 
is  about  7  to  1,  and  taking  the  weight  of  the  aero  motor  as 
one-half  that  of  the  auto  type,  the  true  output  ratio  be- 
comes 14  to  1.  Up  to  the  time  of  a  complete  overhaul  (50 
hours),  and  at  100  miles  per  hour,  the  average  distance 
traveled  by  the  aero  motor  is  5000  miles.  The  equivalent 
motor  car  mileage  is  25,000,  and  the  duration  is  about  1000 
hours.  This  suggests  the  necessity  for  improved  ma- 
terials of  construction.  Even  on  the  present  aeronautic 
motors  the  fiber  stress  in  the  crank-shaft  ranges  from 
120,000  to  140,000  pounds  per  square  inch  against  the 
80,000-pound  stress  used  in  auto  shafts.  The  crank  case 
of  an  aeronautic  motor  must  be  particularly  rigid  to  with- 
stand the  stresses  due  to  the  Hght  mounting,  and  this 
demands  a  higher  grade  metal  than  that  ordinarily  used 
with  automobiles.  Unlike  the  car  engine,  quality  comes 
first  and  price  is  a  secondary  consideration. 

Cooling  Systems.  Both  the  air  and  water  cooling  sys- 
tem is  used,  the  former  for  light  fast  aeroplanes  such  as 
speed  scouts,  and  the  latter  for  the  larger  and  more  heavily 
powered  machines.  Even  in  some  types  of  speed  scouts  the 
air-cooled  motor  has  been  displaced  by  the  water-cooled, 
owing  to  the  fact  that  the  air-cooler  cannot  be  built  satis- 
factorily for  outputs  much  greater  than  110  horsepower. 
By  increasing  the  revolutions  of  the  stationary  water- 
cooled  type  an  increase  in  power  may  be  had  with  the 
same  cylinders,  but  in  the  case  of  the  rotary  air-cooled 
type  the  speed  is  limited  by  the  centrifugal  forces  acting 
on  the  cylinders. 

While  the  weight  of  the  radiator,  water  and  piping 
increase  the  weight  of  the  water-cooled  motor  very  con- 
siderably, the  total  weight  is  not  excessive.  When  the 
fuel  is  considered,  the  total  weight  is  below  that  of  the 


MOTORS 


389 


rotary  when  long  flights  are  attempted.    The  radiator  and 
water  add  complication  and  are  a  source  of  danger.    The 


A  6-CyIinder  Hall-Scott  Motor  Installed  in  a  Martin  Biplane. 

radiators  increase  the  head  resistance  and  add  very  con- 
siderably to  the  maintenance  cost. 

Each  type  of  cooling  has  its  limitations,  and  it  is  hoped 
that  an  improvement  in  cooling  may  be  had  in  the  near 
future.    This  system  should  primarily  reduce  the  size  and 


390 


MOTORS 


resistance  of  the  power  plant,  and  if  possible  the  weight, 
although  the  latter  is  a  secondary  consideration.  At  pres- 
ent the  cooling  system  prevents  even  an  approach  to  the 
true  streamline  form  of  the  body. 

Propeller  Speed.  For  the  best  results,  the  propeller 
speed  should  not  exceed  1200  revolutions  per  minute,  and 
for  structural  reasons  this  is  generally  limited  to  1500 
R.  P.  M.    This  at  once  puts  a  limiting  value  on  the  output 


A  Motor  Installation  in  a  Pusher  Type  Biplane,  Showing  the  Motor  at  the  Rear 
and  the  Double  Radiator  Sections  Over  the  Body. 


of  a  given  size  engme  unless  a  gear  down  arrangement  is 
used.  It  should  be  understood,  between  certain  limits, 
that  the  power  output  increases  roughly  as  the  speed. 
With  direct  drive  arrangements  in  which  the  propeller  is 
mounted  directly  on  the  end  of  the  engine  shaft,  the  motor 
revs,  are  necessarily  the  propeller  revs.,  and  the  only  way 
of  increasing  the  speed  is  by  increasing  the  length  of  the 
stroke  or  by  gearing  down.  An  increase  in  stroke  adds 
rapidly  to  the  weight  by  increasing  the  cylinder  length, 
length  of  connecting  rod,  length  of  crank  throws,  etc. 
Horsepower  Rating.    At  present  there  are  many  meth- 


MOTORS  391 

ods  of  calculating  the  horsepower  of  gasoline  engines. 
Formula  applying  to  auto  or  boat  motors  does  not  apply 
to  flight  conditions,  for  the  aero  motor  is  essentially  a  high 
compression  type  and  has  a  greater  output  per  unit  of 
displacement.  It  is  not  practical  to  give  the  rated  horse- 
power as  the  maximum  output  possible  under  ideal  con- 
ditions, for  this  would  give  no  idea  as  to  the  practical 
capabilities  except  by  long  tedious  calculation.  The  brake 
horsepower  would  give  no  overload  capacity  at  a  fixed 
propeller  speed,  and  the  conditions  are  entirely  different 
from  those  regulating  the  rating  of  auto  motors..  The  lat- 
ter can  be  forced  up  to  the  wrecking  speed,  or  many  times 
the  normal  automobile  speed  of  30  miles  per  hour. 

As  aero  engines  are  generally  well  kept  up,  and  well 
tuned  at  all  times,  the  rated  horsepower  may  be  taken 
from  15  to  20  per  cent  below  that  of  the  maximum  brake 
horsepower.  In  geared-down  motors,  the  gear  efficiency 
is  still  to  be  considered.  The  question  of  the  quality  of 
the  mixture,  and  barometric  pressure,  also  enter  into  the 
problem  whether  the  power  is  rated  on  the  maximum 
obtained  with  a  rich  mixture,  or  is  calculated  from  the 
output  at  the  maximum  efficiency.  A  writer  in  "Aviation" 
suggests  that  the  rated  horsepower  be  taken  as  95  per  cent 
of  the  power  developed  at  a  point  midway  between  the 
maximum  output,  and  the  output  at  the  greatest  efficiency. 
Barometric  pressure  to  be  30  inches  and  the  revolutions 
1200. 

Owing  to  the  great  diversity  in  the  bore-stroke  ratio,  a 
power  formula  must  include  the  bore  and  stroke.  This 
makes  the  S.  A.  E.  formula  for  auto  motors  impossible. 
A  formula  is  proposed  by  a  writer  in  "Aviation."  The 
writer  has  checked  this  up  with  the  published  performance 
of  several  well-known  aeronautical  motors. 

H  =  B-SXR    Where  B  =  bore  in  inches,  S  =  stroke  in 

12.500 


392  MOTORS 

inches,  N  =  number  of  cylinders,  R  =  Crank-shaft  revo- 
lutions per  minute,  and  H  =  rated  horsepower.  This  ap- 
plies only  to  the  four-stroke  cycle  type. 

Power  and  Altitude.  The  power  drops  oft  rapidly  with 
an  increase  in  altitude  unless  corrections  are  made  for 
compression  and  mixture.  With  constant  volume,  the 
decreased  density  causes  decreased  compression.  As  the 
weight  of  air  taken  in  per  stroke  is  reduced,  this  also 
reduces  the  amount  of  fuel  that  can  be  burned  per  stroke. 
By  holding  the  compression  constant  through  adjustment 
of  the  clearance  or  valve  motion,  a  fairly  constant  output 
can  be  had  through  a  wide  range  of  altitudes. 

A  compression  of  115  pounds  per  square  inch  (com- 
monly used)  is  difficult  to  handle  with  a  light  construc- 
tion, but  this  pressure  must  be  obtained  if  the  output  is 
to  be  kept  within  practical  limits.  Engines  having  a  com- 
pression ratio  of  as  high  as  6  are  running  satisfactorily  at 
sea  level,  this  ratio  giving  a  mean  effective  working 
pressure  of  134  pounds  per  square  inch.  With  this  ratio 
the  engine  cannot  be  used  with  full  open  throttle  at  sea 
level  for  more  than  10  or  15  minutes  without  causing  dam- 
age to  the  shaft,  bearing  and  valves.  At  about  10,000  feet 
the  compression  is  normal. 

At  great  altitudes  carburetion  has  become  a  great  prob- 
lem, and  as  aerial  battles  have  already  taken  place  at  ele- 
vations of  20,000  feet,  it  is  quite  possible  that  future 
motors  will  be  equipped  with  some  device  that  will  force  a 
measured  fuel  charge  into  the  cylinders.  The  air  neces- 
sary for  the  combustion  will  also  have  to  be  pumped  in  by 
some  means. 

Weight  Per  Horsepower.  The  weight  per  horsepower 
of  the  engine  is  a  very  loose  term  since  so  much  depends 
upon  the  equipment  included  in  the  weight.  As  many 
as  20  items  may  be  considered  as  being  in  the  doubtful  list, 
and  among  these  are  the  radiator  water,  piping,  mounting, 
propeller  hub,  oil  in  sump,  wiring,  self-starter,  etc.    The 


MOTORS 


393 


only  true  unit  weight  is  that  obtained  by  taking  the  plant 
complete  ( ready  to  run),  with  the  cooling  system,  gasoline 
for  an  hour's  flight,  and  the  oil.  The  weight  of  the  bare 
engine  signifies  nothing.  The  weights  of  the  various  items 
used  on  well  known  motors  are  given  in  a  table  under  the 
chapter  "Weight  Calculations."  While  the  bare  weight  of 
a  certain  engine  may  be  very  low  per  brake  horsepower, 
an  excessive  fuel  consumption  will  often  run  the  effective 
weight  up  and  over  that  of  a  type  in  which  the  bare  weight 
is  far  greater.    The  weight  of  the  engine  per  horsepower. 


Two  Examples  of  Cowls  Used  Over  Rotary  Cylinder  Motors  (Air  Cooled). 

including  the  magneto  and  carbureter,  will  run  from  2.2 
to  5.0  pounds,  according  to  the  type. 

Fuel  Consumption.  The  fuel  consumption  of  water- 
cooled  motors  varies  from  0.48  to  0.65  pounds  per  horse- 
power hour,  an  average  of  0.6  being  safe.  The  fuel  con- 
sumption of  a  rotary  air-cooled  motor  will  range  from  0.6 
to  0.75.  The  oil  consumption  varies  from  0.18  gallons  per 
horsepower  in  the  air-cooled  type  to  0.035  with  the  water- 
cooled  stationary  motor. 

Radiators.  Owing  to  extremes  in  the  temperature  of 
the  air  at  different  altitudes,  the  radiating  surface  should 
be  divided  into  sections  so  that  a  constant  cooling  effect 
can  be  obtained  by  varying  the  effective  surface  of  the 
radiator.    The  temperature  can  also  be  controlled  by  an 


394 


MOTORS 


automatically  regulated  by-pass  which  short  circuits  a 
part  of  the  radiator  water  at  low  temperatures.  Constant 
water  temperature  has  much  to  do  with  the  efficiency  and 
general  operation  of  the  motor,  and  there  will  be  only  one 
temperature  at  which  the  best  results  can  be  obtained. 


^ 

(A)  (B) 

Typical  Radiators.     (A)    Side  or  Top  Type.      (B)   Front  Type. 

Hunsaker  finds  that  0.83  square  feet  of  actual  cooling 
surface  per  horsepower  is  correct  at  60  M.  P.  H.,  while 
others  give  a  value  of  about  1.00  square  foot  under  similar 
conditions.  The  front  or  projected  area  varies  with  the 
thickness  of  the  radiator,  the  thicknesses  varying  from  2 
to  5  inches.  The  Livingston  radiator  gives  a  cooling  sur- 
face of  50  square  inches  per  square  inch  of  front  surface. 
The  total  cooling  effect  depends  upon  the  speed,  the  loca- 


MOTORS  395 

tion  in  regard  to  the  slipstream,  and  the  position  on  the 
body.  A  radiator  maker  should  always  be  consulted  when 
making  the  final  calculations.     See  Chapter  X\'L 

Fuel  Tanks  and  Piping.  The  fuel  tanks  may  be  of  cop- 
per, aluminum  or  tin-coated  steel,  and  all  joints  should  be 
welded  or  riveted.  Never  depend  upon  solder,  as  such 
joints  soon  open  through  the  vibration  of  the  engine. 
Gasoline  should  not  come  into  contact  with  steel,  nor  the 
zinc  used  on  galvanized  iron.  Splash  plates  are  provided 
to  keep  the  fluid  from  surging  back  and  forth  while  in 
flight.  All  gas  should  be  supplied  to  the  engine  through 
a  filter  or  strainer  placed  in  the  main  gas  line.  The  valves 
in  the  fuel  lines  should  be  provided  with  stopcocks,  so 
arranged  that  they  can  be  closed  from  the  pilot's  seat. 

In  general,  the  carbureters  should  be  fed  by  gravity 
from  an  overhead  service  tank,  this  tank  being  suppHed 
from  the  main  reservoir  by  air  pressure  or  a  gasoline 
pump.  The  air  can  be  compressed  by  a  pump  on  the 
engine  or  by  a  paddle  driven  pump  operated  by  the  air- 
stream,  and  as  a  rule  the  latter  is  preferable,  as  it  can  be 
operated  with  the  aeroplane  gliding  and  with  the  engine 
dead.  Air  pressure  systems  are  likely  to  fail  through 
leaks,  while  with  a  good  gasoline  pump  conditions  are 
much  more  positive.  The  gravity  service  tank  should  be 
located  so  that  it  will  feed  correctly  with  the  aeroplane 
tilted  at  least  30  degrees  from  the  horizontal. 

The  gasoline  piping  should  be  at  least  5/16  inch  inside 
diameter,  and  should  be  most  securely  connected  and  sup- 
ported against  vibration.  To  guard  against  crystallization 
at  the  point  of  attachment,  special  flexible  rubber  hose  is 
generally  used.  This  must  be  hose  made  specially  for 
this  purpose,  as  ordinary  rubber  hose  is  soon  dissolved  or 
rotted  by  gasoline  and  oil.  Air  pockets  must  be  avoided  at 
every  point  in  the  fuel- and  oil  system.    ' 

Rotating  Cylinder  Motors.  The  first  rotating  cylinder 
motor  in  use  was  the  American  Adams-Farwell,  a  type 


396 


MOTORS 


that  Avas  soon  followed  by  the  better  known  French 
"Gnome."  Other  motors  of  this  type  are  the  Clerget, 
LeRhone,  Gyro  and  Obereusel.  They  are  all  of  the  air- 
cooled  type — cooled  partly  by  the  revolution  of  the  cyl- 
inders about  the  crank-shaft,  and  partly  by  the  propeller 
slipstream.  While  the  pistons  slide  through  the  cylinder 
bore,  the  rotating  cylinder  motor  is  not  truly  a  recip- 
rocating type,  as  the  pistons  do  not  move  back  and  forth 


Two  Views  ot  the  "Monosoupape"  Gnome  Rotary  Cylinder  Motor.  This  Motor 
Has  9  Cylinders  Arranged  Radially  Around  the  Crankshaft  and  Develops 
100  Horsepower.    The  Cylinders  Are  Air  Cooled. 


in  regard  to  the  crank  shaft.  The  cylinders  revolve  about 
the  crank  shaft  as  a  center,  while  the  pistons  and  connect- 
ing rods  revolve  about  the  crank  pin,  the  difference  in 
the  pivot  point  causing  relative,  but  not  actual,  recip- 
rocation. 

The  original  Gnome  motor  drew  in  the  charge  through 
an  inlet  valve  in  the  piston  head.  The  gas  passed  from 
the  mixer,  through  the  hollow  crank-shaft,  and  then  into 
the  crank-case.  The  exhaust  valve  was  in  the  cylinder 
head.  This  valve  arrangement  was  not  entirely  satis- 
factory, and  the  company  developed  the  "Monosoupape" 


MOTORS 


39' 


Hall-Scott  "Big  Six"  Aeronautical  Motor  of  the  Vertical  Watercooled  Type. 
125  Horsepower. 


Hall-Scott  4-CyHnder  Vertical  Watercooled  Motor.     80-90  Horsepower. 


398 


MOTORS 


or  "Single  valve"  type.  The  100  H.  P.  Monosoupape 
Gnome  has  9  cylinders,  4.3"  x  5.9".  The  total  weight  is 
272  pounds  and  the  unit  weight  is  2.72  pounds  per  horse- 
power. It  operates  on  the  four-stroke  cycle  principle.  The 
gas  consumption  is  12  gals,  per  hour,  and  it  uses  2.4  gals, 
of  castor  oil.  The  cylinders  and  cooling  fins  are  machined 
from  a  solid  steel  forging,  weighing  88  pounds.  The 
finished  cylinder  weighs  5.5  pounds  after  machining.    The 


Sturtevant  "V"  Type  8-Cylinder  Water  Cooled  Aeronautical  Motor.    This  Motor 
Is  Provided  With  a  Reduction  Gear  Shown  at  the  Rear  of  the  Crankcase. 

walls  are  very  thin,  probably  about  1/16  inch,  but  they 
stand  up  well  under  service  conditions. 

Assuming  the  piston  to  be  on  the  compression  stroke, 
the  ignition  will  occur  from  15°  to  20°  before  the  top  dead 
center.  Moving  down  on  the  working  stroke,  and  at  85° 
from  top  dead  center,  the  exhaust  valve  begins  to  open, 
and  the  exhaust  continues  until  the  piston  returns  to  the 
upper  dead  center.  With  the  valve  still  open,  pure  air 
now  begins  to  enter  through  the  exhaust  valve  and  con- 
tinues to  flow  until  the  valve  closes  at  65°  below  the  bot- 


AIOTORS 


39f) 


torn  center.  Still  descending,  the  piston  forms  a  partial 
vacuum  in  the  cylinder,  until  at  2°  before  the  lower  cen- 
ter the  piston  opens  the  ports  and  a  very  rich  mixture  is 
drawn  in  from  the  crank  case.  This  rich  mixture  is 
diluted  to  the  proper  density  by  the  air  already  in  the 
cylinder,  and  forms  a  combustible  gas.  The  upward 
movement  of  the  piston  on  the  compression  stroke  closes 


uction  Gear,     Four 
alve  Motion. 


Dusenberg  4-Cylinder  Vertical  Water  Cooled  Motor  With  Red 
valves  Are  Used  Per  Cylinder.     Note  Peculiar  Valvi 

the  ports  and  compression  begins.  The  mixture  enters 
the  crank  case  through  a  hollow  shaft,  with  the  fuel  jets 
near  the  crank  throws.  A  timed  fuel  pump  injects  the  fuel 
at  the  proper  intervals. 

Curtiss  Motors.  The  Curtiss  motors  are  of  the  water- 
cooled  "V"  type,  with  6  to  8  cylinders  per  row.  These  are 
probably  the  best  known  motors  in  America  and  are  the 
result  of  years  of  development,  as  Curtiss  was  the  first 
to  manufacture  aero  motors  on  a  practical  scale. 


400 


MOTORS 


Hall-Scott  Motors.  These  motors  are  made  by  one  of 
the  pioneer  aeronautical  motor  builders,  and  have  met 
with  great  favor.  They  are  of  the  vertical  water-cooled 
type,  and  with  the  exception  of  minor  details  and  weight 
are  very  similar  in  external  appearance  to  the  automobile 
motor.    Four  and  6-cylinder  types  are  built. 


Roberts  Two-Stroke  Cycle  Aeronautical  Motor,  of  the  Vertical  6-Cylinder  Type. 
Two  Carbureters  Are  Used,  One  for  Each  Group  of  Three  Cylinders. 


Sturtevant  Motors.  These  are  of  the  "V"  water-cooled 
type,  and  are  provided  with  or  without  a  reducing  gear. 
At  least  one  model  is  provided  with  lined  aluminum  cyl- 
inders. 

Dusenberg  Motor.  This  is  a  four-cylinder,  water- 
cooled,  vertical  motor  with  a  very  peculiar  valve  motion. 
The  valves  are  operated  by  long  levers  extending  from 
the  camshaft.  Two  inlet,  and  two  exhaust  valves,  are  used 


MOTORS  401 

per  cylinder.  The  motor  is  generally  furnished  with  a 
reducing  gear. 

Roberts  Motor.  This  is  a  solitary  example  of  the  two- 
stroke  cycle  type,  and  has  been  used  for  many  years.  It 
is  simple  and  compact,  and  is  noteworthy  for  the  simplicity 
of  its  oiling  system.  The  oil  is  mixed  with  the  gasoline, 
and  is  fed  through  the  carbureter.  This  is  one  of  the  many 
advantages  of  a  two-stroke  cycle  motor. 

Table  of  Aeronautical  Motors.  The  following  table  will 
give  an  idea  as  to  the  general  dimensions  of  American 
aeronautical  motors : 


i|  1^    ^^    £      £       sS     &£     Is     £l  §s  I& 

^S                                       SiS  ZU      m  W3  CSJCrS  pHtt5  PS  P5^  JO      ^O' 

Aeromarine    85       6     4.3"  5.6"  1400  1150  449  440  58"     30" 

Aeromarine    (D-12)...160  12* 1400  750  67"     24" 

Curtiss     OX-2 85       8*4.0"  5.0"  1400  1400      375  50"     30" 

Curtiss   OXX-2 100  8*4.3"  5.0"  1400  1400  568  423  50"     30" 

Curtiss    VX 160  8*5.0"  7.0"  1400  1400  1100  645  68"     35" 

Curtiss  VX-3 200       8*  5.0"  7.0"  1100  667      

Curtiss    V-4 250  12*5.0"  7.0"  1650  1125  85"     35" 

Dusenberg    140  4     5.0"  7.0"  2100  496  455  44"     16" 

Gnome     100       9     4.3"  5.9"  1200  1200  920  272     

Gvro    (K) 90       7     4.5"  6.0"  1250  1250     242      

Gvro    (L) 100       9     4.5"  6.0"  1200  1200  859  285      

Hall-Scott     (A-7) 80-90  4     5.0"  7.0"  1370  1370  550  410  40"     19" 

Hall-Scott     (A-5) 125  6     5.0"  7.0"  1300  1300  825  592  64"     19" 

Hall-Scott     (A-5-a) 162       6     1325  1325      562     

Hispano-Suiza    154  8*4.6"  5.0"  1500  672  455  53"     33" 

Knox    300  12*4.7"  7.0"  1800  1555  1425      

Sturtevant    (5) 140  8*4.0"  5.5"  2000  522  600  60"     23"" 

Sturtevant   (5-a) 140  8*4.0"  5.5"  2000  522  514  60"      ... 

Motors  marked  (*)  are  of  the  "V"  type.  ,«ao  JT«»  >*i«'  i*a       2.i^f 

The  Liberty  Motor.  The  necessity  of  speed  and  quan- 
tity in  the  production  of  aeronautical  motors  after  the 
declaration  of  war  caused  the  Government  to  seriously 
consider  the  design  of  a  highly  standardized  motor.  This 
idea  was  further  developed  in  a  conference  with  repre- 
sentatives of  the  French  and  British  missions  on  May 
28,  1917,  and  was  then  submitted  in  the  form  of  sketches 
at  a  joint  meeting  of  our  allies,  the  Aircraft  Production 
Board,  and  the  Joint  Army  and  Navy  Technical  Board. 
The  speed  with  which  the  work  was  pushed  is  remarka- 
ble, for  on  July  3rd,  the  first  model  of  the  eight  cylinder 


402  MOTORS 

type  was  delivered  to  the  Bureau  of  Standards.  Work 
was  then  concentrated  on  the  12  cyHnder  model,  and  one 
of  the  experimental  engines  passed  the  50  hour  test 
August  25,  1917. 

It  is  of  the  "V"  type  with  the  cylinder  blocks  at  an 
angle  of  45  degrees  instead  of  60  degrees  as  in  the  ma- 
jority of  12  cylinder  "V"  motors.  This  makes  the  motor 
much  narrower  and  more  suitable  for  installation  in  the 
fuselage,  and  in  this  respect  is  similar  to  the  arrangement 
of  the  old  Packard  aviation  motor.  It  has  the  additional 
advantages  of  strengthening  the  crank  case.  The  bore 
and  stroke  is  5''x7"  as  in  the  Hall-Scott  models  A-5  and 
A-7.  The  cylinders  combine  the  leading  features  of  the 
German  Mercedes,  the  English  Rolls-Royce,  Lorraine- 
Dietrich,  and  Isotta-Fraschini.  Steel  cylinder  walls  are 
used  with  pressed  steel  water  jackets,  the  latter  being 
applied  by  means  of  a  method  developed  by  the  Packard 
Company.  The  valve  cages  are  drop  forgings,  welded 
to  the  cylinder  heads. 

The  camshaft  and  valve  gear  are  above  the  cylinder 
head  as  in  the  Mercedes,  but  the  lubrication  of  the  parts 
was  improved  upon  by  the  Packard  Company. 

The  crankshaft  follows  standard  12  cylinder  practice 
except  as  to  the  oiling  system,  the  latter  following  Ger- 
man practice  rather  closely.  The  first  system  used  one 
pump  to  keep  the  crankcase  empty  delivering  the  oil  to  an 
outside  reservoir.  A  second  pump  took  the  oil  from  the 
reservoir  and  delivered  it  to  the  main  crankshaft  bearings 
under  pressure.  The  overflow  from  the  main  bearings 
traveled  out  over  the  face  of  the  crank  throw  cheeks  to 
a  "Scupper,"  which  collected  the  excess  for  crank  pin 
lubrication.  In  the  present  system,  a  similar  general 
method  is  followed  except  that  the  pressure  oil  is  not 
only  fed  to  the  main  crankshaft  bearings,  but  also  through 
holes  in  the  crank  cheeks  to  the  crank  pins  instead  of  by 
the  former  scupper  feed. 


MOTORS 


403 


A  special  Zenith  carburetor  is  used,  that  is  particularly 
adapted  to  the  Liberty  motor.  A  Delco  iguitiou  system  of 
special  form  is  installed  to  meet  the  peculiar  cylinder  block 
angle  of  45  degrees.  This  ignition  is  of  the  electric  gen- 
erator type  and  magnetos  are  not  used. 

Several  American  records  have  been  broken  by  the  new 
motor,  and  it  is  reported  to  have  given  very  satisfactory 
service,  but  full  details  of  the  performance  are  difficult  to 
obtain  owing  to  the  strict  censorship  maintained  in  regard 
to  things  aeronautic.  The  motor  is  particularly  well 
adapted  to  heavy  bombing  and  reconnaissance  type  ma- 
chines, or  for  heavy  duty.  It  is  reported  that  the  use  of 
the  motor  has  been  discontinued  on  speed  scouts,  although 
further  developments  along  this  line  may  not  have  been 
reported. 

The  following  gives  the  principal  characteristics  of  the 
Liberty  motor,  issued  by  the  National  Advisory  Committee 
for  Aeronautics. 


Gasoline 

Year 

Horse- 

Weight 

Weight 

H.P. 

(Model) 

power 

Pounds 

Per  H.  P. 

Hour 

1917 

400 

801 

2.00 

0.50 

1918 

432 

808 

1.90 

0.48 

1918 

450 

825 

1.80 

0.46 

The  motors  listed  are  all  12  cjdinder  models,  and  the  out- 
put and  unit  weights  are  based  on  a  crank-shaft  speed  of 
1800  R.  P.  :\L  The  5"x7"  bore  and  stroke  give  an  output 
of  37.5  horsepower  per  cylinder  in  the  latest  model.  In 
1917,  the  Liberty  motor  was  65  per  cent  more  powerful, 
and  28  per  cent  lighter,  than  the  average  stock  motor  in 
service  during  that  year. 


CHAPTER  XXI. 
GLOSSARY  OF  AERONAUTICAL  WORDS. 

In  the  following  list  are  the  most  common  of  the  aeronautical 
words  and  phrases.  Many  of  these  words  are  of  French  origin, 
and  in  such  cases  are  marked  "Fr."  In  cases  of  English  words, 
the  French  equivalents  follow  in  parentheses.  When  a  French  word 
or  term  is  given  it  is  in  italics,  unless  it  is  in  common  use  in  this 
country.  Words  marked  (*)  are  the  revisions  adopted  by  the  Na- 
tional Advisory  Board  of  Aeronautics  at  Washington,  D.  C,  and 
include  the  term  "Airplane,"  which  was  intended  to  supplant  the 
more  common  "Aeroplane."  These  revisions  have  not  met  with 
universal  adoption,  for  the  older  words  are  too  well  established  to 
admit  of  change. 

A 

ABSOLUTE  UNITS.  Units  given  in  terms  of  mass.  For  expres- 
sion in  terms  of  pounds  (Gravitational  units)  they  must  be  multi- 
plied by  some  factor  involving  the  value  of  gravitation.  Thus, 
to  convert  units  of  mass  into  pounds,  the  mass  must  be  multiplied 
by  the  value  of  gravitation,  32.16  being  the  average  figure  taken 
for  this  quantity.  To  convert  the  absolute  lift  factors  given  by 
the  N.  P.  L.  into  pounds  per  square  foot  per  mile  per  hour,  mul- 
tiply the  absolute  value  by  0.0051V". 

ABSOLUTE  ZERO.  The  temperature  at  which  heat  ceases  to 
exist.  This  is  461  degrees  below  the  Fahrenheit  zero,  or  273  de- 
grees below  the  Centigrade  zero. 

ACCELERATE.    To  increase  in  speed. 

ACIER  (Fr.)     Steel. 

AERODONETICS.  A  word  originated  by  Lanchester  to  denote 
the  science  of  stability. 

AEROCURVE.     See  AEROFOIL. 

AEROFOIL.*  A  thin  wing-like  structure  designed  to  obtain  lift 
by  the  reaction  of  moving  air  upon  its  surfaces. 

AERODYNAMICS.  A  science  investigating  the  forces  produced 
by  a  stream  of  air  acting  upon  a  surface. 

AERODYNAMIC  RESISTANCE.  The  resistance  caused  by  tur- 
bulence or  eddies. 


GLOSSARY 

AERODROME.    A  flying  field.    This  word  was  also  used  by  Lang- 
ley  to  describe  an  aeroplane. 
AEROSTAT.    A  lighter  than  air  machine. 
AEROSTATICS.     The   science   of   lighter   than   air   machines,   or 

devices  sustained  by  flotation. 
AEROPLANE.     (Fr.  L'Avion.)     A  heavier  than  air  craft  sustained 

by  fixed  wing  surfaces  driven  through  the  air  at  the  same  velocity 

as  the  body  of  the  machine.    Auxiliary  surfaces  are  provided  for 

stabilizing,  steering,   and    for   producing  changes   in   the   altitude. 

The  landing  gear  may  be  suitable  for  either  land  or  water,  although 

in  the  latter  case   it   is  generally  known   as  a   "Seaplane."     The 

Committee  equivalent  is  "Airplane." 
AILERON.*     A   movable   auxiliary   surface   used    in    maintaining 

lateral  balance. 
AILE  (Fr.)     Wing. 
AIRBOAT.    An  aeroplane  provided  with  a  light  boat  hull  in  which 

the  pilot  and  passenger  are  enclosed. 
AIRCRAFT.*     Any  form  of  craft  designed  for  the  navigation  of 

the  air.    This  includes  aeroplanes,  balloons,  dirigibles,  helicopters, 

ornithopters.  etc. 
AIRPLANE.*     See  aeroplane. 

AIRSHIP.     A  lighter  than  air  craft  provided  with  means  of  pro- 
pulsion. 
AIR-SCREW.    See  PROPELLER. 
ALTIMETER.*    An  instrument  used  for  determining  the  height  of 

aircraft  above  the  earth. 
ALTITUDE.     Height  of  aircraft  above  sea  level— generally  given 

in  feet. 
AMPHIBIAN.    An  aeroplane  equipped  with  landing  gear  for  both 

land  and  water. 
ANEMOMETER.*     An  instrument  for  measuring  the  velocity  of 

the  wind. 
ANGLE  OF  INCIDENCE.    The  angle  made  by  a  surface  or  body 

with  an  air  stream.     In  the  case  of  curved  wings,  the  angle  is 

measured  from  the  chord  of  the  curve. 
ANGLE  OF  ATTACK.*     See  Angle  of  Incidence. 
ANGLE  OF  ENTRY.     The  angle  made  with  the  chord  of  a  wing 

section  by  a  line  drawn  tangent  to  the  upper  curved  face,  and  at 

the  front  edge. 
ANGLE  OF  TRAIL.     The  angle  made  by  a  line  drawn  tangent  to 

the  upper  surface  at  the  trailing  edge. 
ANEMOGRAPH.     An  instrument  used  for  graphically  recording 

the  velocity  of  air  currents. 


GLOSSARY 

APTEROID  ASPECT.  A  wing  is  in  apteroid  aspect  when  the 
narrow  edge  is  toward  the  wind. 

ARETIER  ARRIERE  (Fr.)     See  Trailing  Edge. 

ARETIER  AVANT  (Fr.)     See  Leading  Edge. 

ARBRE  (Fr.)     Shaft. 

ASPECT  RATIO.*  The  ratio  of  the  wing  span  to  the  chord 
(length  divided  by  width). 

ATTERRISSAGE   (Fr.)     Landing  Gear. 

ATTACHES  (Fr.)     Fastenings. 

AUXILIARY  SURFACE.  A  surface  used  for  stability  or  for  the 
control  of  the  aeroplane. 

AVIAPHONE.  An  electric  system  of  communication  between  the 
passenger  and  pilot. 

AVION  (Fr.)     See  Aeroplane. 

AXIS  OF  PITCH.  The  axis  taken  parallel  to  the  length  of  the 
wings,  and  through  the  center  of  gravity.  This  is  sometimes  called 
the  "Y"  axis. 

AXIS  OF  ROLL.  An  axis  passing  fore  and  aft  parallel  to  the 
center  line  of  the  propeller.  This  axis  is  sometimes  called  the 
"X"  axis. 

AXIS  OF  YAW.  A  vertical  axis,  passing  through  the  center  of 
gravity,  around  which  the  machine  swings  when  being  steered  in 
a  horizontal  direction  under  the  action  of  the  rudder.  This  is  the 
"Z"  axis. 

B 

BALSA  WOOD.  A  very  light  wood  obtained  from  South  America. 
It  is  lighter  than  cork. 

BALLOON.*  A  form  of  aircraft  of  the  lighter  than  air  type  com- 
prising a  gas  bag  and  car.  It  is  not  provided  with  a  power  plant, 
and  depends  upon  the  bouyancy  of  the  gas  for  its  sustentation. 
A  balloon  restrained  from  free  flight  by  means  of  a  cable  is 
known  as  a  "Captive  balloon."  A  kite  balloon  is  an  elongated 
form  of  captive  balloon,  fitted  with  a  tail  to  keep  it  headed  into 
the  wind,  and  is  inclined  at  an  angle  so  that  the  wind  aids  in 
increasing  the  lift  of  the  gas. 

BALLONET.*  A  small  air  balloon  within  the  main  gas  bag  of  a 
balloon  or  dirigible  used  for  controlling  the  ascent  or  descent,  and 
for  keeping  the  fabric  of  the  outer  envelope  taut  when  the  pres- 
sure of  the  gas  is  reduced.  The  ballonet  is  kept  inflated  with  air 
at  the  required  pressure,  the  air  being  controlled  by  a  valve  or  by 
regulating  the  speed  of  the  blower. 

BANK.*  To  incline  the  wings  laterally  when  making  a  turn  so  that 
a  portion  of  the  lift  force  will  be  opposed  to  the  centrifugal  force. 


GLOSSARY 

BAROGRAPH.*  (Fr.  Barographe.)  An  instrument  used  for  re- 
cording pressure  variations  in  the  atmosphere.  The  paper  charts 
on  which  the  records  are  made  are  used  for  determining  the  alti- 
tude of  aircraft. 

BAROMETER.  An  instrument  used  for  measuring  variations  in 
the  atmospheric  pressure,  but  is  not  provided  with  a  recording 
mechanism  as  in  the  case  of  the  barograph. 

BEAUME.  A  scale  of  density  or  a  hydrometer  unit  used  in  meas- 
uring the  density  of  fluids.  On  the  Beaume  scale  water  is  lo.oo, 
while  on  the  "Specific  gravity"  scale  water  is  i.oo.  The  Beaume 
scale  is  generally  used  for  gasoline  and  oils. 

BENDIXG-MOMENT.  The  moment  or  "Leverage"  that  tends  to 
bend  a  beam. 

BEQUILLE   (Fr.)     Tail  Skid. 

BERCEAU  de  MOTEUR  (Fr.)     Engine  Bed. 

BIAS  LAID  FABRIC.  Fabric  laid  on  the  wing  structure  with  the 
seams  at  an  angle  with  the  ribs. 

BIPLACE   (Fr.)     Two  Seater. 

BIPLANE.*  (Fr.  Biplan.)  An  aeroplane  with  two  superposed 
lifting  surfaces. 

BODY.    See  FUSELAGE. 

BODY  RAILS.     See  LONGERONS. 

BO  IS  (Fr.)     Wood. 

BOIS  CREUS  (Fr.)     Hollow  wood  construction. 

BOMBER.    An  aeroplane  used  for  bombing  operations. 

BOOM.  The  fore  and  aft  beams  running  from  the  wings  to  the 
tail  in  a  pusher  type  biplane. 

BORD  de  ATT  AC  QUE  (Fr.)     Entering  or  leading  edge. 

BORD  de  SORTIE  (Fr.)     Trailing  edge. 

BLADE,  PROPELLER.     (Fr.  Pale-Helice.) 

BOULON  (Fr.)     Bolt. 

BOUSSOLE  (Fr.)     Compass. 

BRAS  de  AILE  (Fr.)     Wing  Spar. 

BREVET    (Fr.)     Flying  permit  or  license. 

BRAKES,  AIR.  Small  adjustable  flaps  used  in  increasing  the  head 
resistance  during  a  landing,  thus  decreasing  the  speed. 

BURBLE  POINT.  The  angle  at  which  the  lift  of  a  wing  section 
reaches  a  maximum. 

BUOYANCY.  The  static  force  due  to  a  difference  in  density.  The 
difference  in  density  between  the  gas  in  a  balloon  envelope  and  the 
outside  air  determines  the  sustaining  or  buoyant  force  of  a  balloon. 

BUS.    A  slow  fairly  stable  aeroplane  used  in  training  schools. 


GLOSSARY 


CABLE.  (Fr.  Cable.)  A  wire  rope  built  up  of  a  number  of  small 
strands. 

CAERE' *  (Fr.)  A  %ing  attitude  in  which  the  angle  of  incidence 
is  larger  than  normal  with  the  tail  well  down. 

CAMBER.*  The  convexity,  or  rise  of  a  lifting  surface,  measured 
from  the  chord  of  the  curve.  It  is  usually,  given  as  the  ratio  of 
the  maximum  height  of  the  curve  to  the  length  of  the  chord.  Top 
camber  refers  to  the  upper  surface,  and  bottom  camber  to  the 
lower  surface. 

CABANE  (Fr.)  The  center  struts  rising  from  the  top  of  the  body 
to  the  upper  wing,  or  the  short  struts  used  for  the  bracing  of  the 
overhanging  portions  of  a  biplane  wing.  Usually  cabane  denotes 
the  center  cell  struts. 

CANARD  (Fr.)  A  machine  in  which  the  elevator  and  stabilizer 
are  in  front.     The  canard  type  flies  "Tail  first." 

CAPOT  (Fr.)     Cowl  or  motor  hood. 

CAPTIVE  BALLOON.*     See  Balloon. 

CAPACITY.*  The  lifting  capacity  is  the  maximum  flying  load  of 
an  aicraft.  The  carrying  capacity  (live  load)  is  the  excess  of  the 
lifting  capacity  over  the  dead  weight  of  the  aeroplane,  the  latter 
including  the  structure,  power  plant  and  essential  accessories. 

CARLINGUE  (Fr.)     Cock-pit. 

CATA-HEDRAL.  A  negative  dihedral,  or  wing  arrangement,  where 
the  wing  tips  are  lower  than  the  center  portion. 

CATAPULT.  A  device  for  launching  an  aeroplane  from  the  deck 
of  a  ship  or  other  limited  space.  The  first  Wright  machines  were 
launched  with  a  catapult. 

CELL.  (Fr.  Cellule.)  The  space  included  between  adjacent  struts 
of  a  biplane.  The  space  between  the  center  struts  is  the  "Center 
Cell." 

CEILING.  The  maximum  altitude  to  which  an  aeroplane  can 
ascend. 

CEINTURE  de  SURETE  (Fr.)     Safety  Belt. 

CENTER  OF  PRESSURE  (C.  P.)*  The  point  of  application  of 
the  resultant  of  all  aerodynamic  forces  on  an  aeroplane  wing.  If 
the  wing  is  supported  at  the  center  of  pressure  it  will  be  in  equi- 
librium. 

CENTER  OF  GRAVITY.  The  point  at  which  an  aeroplane  will 
balance  when  freely  suspended. 

CENTER  OF  BUOYANCY.*  The  point  at  which  the  resultant 
of  all  the  buoyant  forces  act. 


GLOSSARY 

CHARXIERE  (Fr.)     Hinge. 

CHASSIS.  The  landing  wheels  and  their  frame.  This  is  also  called 
the  "Landing  gear"  in  English,  or  the  "Train  de  Atferrissage" 
in  French.  The  chassis  carries  the  load  when  resting  on  the 
ground  or  when  running  over  the  surface. 

CHORD.*  This  has  two  meanings.  The  chord  is  the  width  of  a 
wing  or  its  shortest  dimension.  The  chord  is  also  the  straight  line 
drawn  across  the  leading  and  trailing  edges  of  a  wing  section. 

CHASER.  (Fr.  Avion  de  Chasse.)  A  small,  fast  machine  used  in 
scouting  or  fighting.     This  type  is  also  known  as  a  "Speed  scout." 

CHARA-A-BANC  (Fr.)  A  two  seater  aeroplane  in  which  the  pilot 
and  passenger  are  seated  side  by  side. 

CLOCHE  (Fr.)  A  type  of  control  column  used  on  the  old  Type  XI 
Bleriot. 

COCK-PIT.  The  part  of  the  body  occupied  by  the  pilot  or  passen- 
ger. The  openings  in  the  body  cut  for  entrance  and  exit  are  the 
"Cock-pit  Openings." 

COMMAXDES  A  POXT  (Fr.)  Control  bridge  or  Deperdussin 
yoke. 

COMPTE  TOURS  (Fr.)    Tachometer  or  speed  indicator. 

COMPOXEXTS.  The  individual  forces  that  make  up  a  total  re- 
sultant force. 

COXTROLS.*  (Fr.  Commandes.)  The  complete  system  used  for 
steering,  elevating,  balancing,  and  speed  regulation.  When  con- 
trols are  operated  by  hand  they  are  known  as  "Manual  Controls." 

COXTROL  BRIDGE.  (Fr.  Commandes  A  Pont.)  The  "U"  shaped 
lever  used  with  the  Deperdussin  control  system.  Sometimes  known 
as  the  "Yoke." 

COXTROL  STICK.  (Fr.  Manche  A  Balqi.)  A  simple  control 
lever  capable  of  being  moved  in  four  directions  for  elevation, 
depression  and  lateral  balance. 

COXTROL  SURFACES.  The  adjustable  surfaces  used  for  direct- 
ing and  balancing  aircraft.  On  an  aeroplane  these  are  represented 
by  the  rudder,  elevator,  and  ailerons. 

COXTREPLAQVE  (Fr.)     Three-ply  wood. 

CORDE  (Fr.)     Cord  or  wire. 

CORD  A  PIAXO  (Fr.)     Piano  or  solid  hard  wire. 

CORD  WIXDIXG  (Fr.  Transfil.)  A  winding  wrapped  around 
wooden  struts  to  prevent  splintering  or  complete  fracture. 

COSSE  (Fr.)     Thimble  for  cable  connections. 

COUSSIX  (Fr.)     Cushion. 

COVERIXG,  WIXG.  (Fr.  Entoilage.)  The  fabric  used  in  cover- 
ing the  wing  structure. 


GLOSSARY 

CRITICAL  ANGLE.*  The  angle  of  attack  or  incidence  at  which 
the  lift  is  a  maximum. 

COWL.  (Fr.  Capot.)  The  metal  cover  surrounding  a  rotary  cyl- 
inder motor. 

CROISILLONS  (xV.)     Bracing  wires. 


DAMPING.  The  reduction  of  oscillation  or  vibration  by  the  resist- 
ance of  the  stabilizing  surfaces. 

DEAD  LOAD.  The  weight  of  the  structure,  power  plant,  and  essen- 
tial accessories. 

DEAD  WATER.  The  wake  directly  in  the  rear  of  a  moving  body  or 
surface. 

Dc  CHASSE  (Fr.)     See  CHASER. 

DECALAGE.*  The  difference  in  the  angle  of  incidence  between 
the  upper  and  lower  wings  of  a  biplane. 

DEMOISELLE  TYPE.  A  small  monoplane  type  developed  by 
Santos  Dumont. 

DENSITY.    The  specific  weight,  or  the  weight  per  cubic  foot. 

DIEDRE  (Fr.)     Dihedral  angle. 

DERIVE  (Fr.)     Fin. 

DIHEDRAL  ANGLE.  (Fr.  Diedre.)  When  the  tips  of  the  wing 
are  higher  than  at  the  center,  the  two  wing  halves  form  an  angle. 
The  included  angle  between  the  two  halves,  taken  above  the  sur- 
face, is  known  as  the  "Dihedral  angle." 

DIPPING  FRONT  EDGE.  A  wing  section  in  which  the  leading 
edge  is  well  bent  down  below  the  rest  of  the  lower  surface. 

DIRIGIBLE.*  A  lighter  than  air  craft  in  which  sustentation  is 
provided  by  a  gas  bag.  It  differs  from  a  balloon  in  having  a  power 
plant,  and  is  thus  capable  of  flying  in  any  desired  direction  regard- 
less of  the  wind. 

DISC  AREA  OF  PROPELLER.*  The  total  area  of  the  disc  swept 
out  by  the  propeller  tips. 

DISCONTINUITY.  Interruption  in  direction,  or  breaks  in  stream 
line  flow.  A  body  causing  eddies  or  turbulence  causes  "Discontinu- 
ous flow."  The  surface  separating  the  eddies  and  the  continuous 
stream  is  called  a  "Surface  of  Discontinuity." 

DISPLACEMENT.  The  volume  or  space  occupied  by  a  floating 
body. 

DOUBLE  CAMBER.  A  wing  section  in  which  both  the  top  and 
bottom  surfaces  are  given  a  convex  camber  or  curvature. 

DOPE.  (Fr.  Enduit.)  A  solution  used  for  protecting  and  stretch- 
ing the  wing  fabric. 


GLOSSARY 

DRAG.  The  resistance  offered  to  the  forward  motion  of  a  surface 
or  body  moving  through  the  air.  As  defined  by  the  Advisory 
Committee  this  is  the  total  resistance  offered  by  the  craft  and 
includes  both  the  resistance  of  the  wings  and  body.  This  concep- 
tion is  confusing,  hence  the  author  has  considered  drag  as  being 
the  forward  resistance  of  the  wings  alone.  The  resistance  of  the 
structure  is  simply  called  the  "Head  resistance,"  and  the  sum  of 
the  resistances  is  the  "Total  resistance."  This  nomenclature  was 
in  existence  before  the  Advisory  Board  proposed  their  definition. 

DRIFT.  As  defined  by  the  Advisory  Board,  the  drift  is  the  hori- 
zontal resistance  offered  by  the  wings  alone.  This  is  confusing 
since  previous  works  defined  "Drift"  as  the  amount  by  which  an 
aircraft  was  driven  out  of  its  normal  path  by  wind  gusts.  Accord- 
ing to  usage,  "Drift"  is  the  sidewise  deviation  from  the  normal 
flight  path. 

DRAG  WIRES.  The  bracing  wires  used  for  resisting  the  drag 
stresses  set  up  in  the  wing. 

DRIFT  INDICATOR.  An  instrument  for  indicating  the  amount 
by  which  an  aircraft  is  blown  out  of  its  path  by  side  winds. 

DUAL  CONTROL.  A  double  system  of  control  that  can  be  operated 
both  by  the  pilot  and  passenger. 

DUTCH  ROLL.  A  combined  side  roll  and  fore  and  aft  pitch.  The 
machine  rolls  from  side  to  side  in  combination  with  an  up  and 
down  motion  of  the  nose. 

DYNAMIC  PRESSURE.  The  pressure  due  to  the  impact  of  an 
air  stream. 


ECCENTRIC  LOAD.  A  load  acting  to  one  side  of  the  center  line 
of  a  beam  or  strut. 

ECOLE  (Fr.)     School. 

ECROU   (Fr.)      Nut. 

EDDY.  An  irregularly  moving  mass  of  air  caused  by  the  breaking 
up  of  a  continuous  air  stream,  or  by  "Discontinuity." 

EFFICIENCY.  The  efficiency  of  a  lifting  surface  is  generally 
expressed  by  the  ratio  of  the  lift  to  the  drag,  or  the  "Lift-drag 
ratio."  The  efficiency  of  a  propeller  is  the  ratio  of  the  work 
usefully  applied  to  the  air  stream  in  regard  to  the  power  supplied 
to  the  propeller. 

ELEVATOR.*  The  hinged  horizontal  tail  surface  used  for  main- 
taining longitudinal  equilibrium  and  for  ascent  or  descent. 

EMPENNAGE  (Fr.)  The  group  of  tail  surfaces,  including  the 
elevator  and  stabilizer. 


GLOSSARY 

ENDUIT  (Fr.)     Dope. 

ENGINE   ROTATION.      According   to   the   Advisory   Board,    an 

engine  is  turning  in  right-hand  rotation  when  the  output  shaft  stub 

fs  turning  anti-clockwise. 
ENGINE  BEARERS    (BED).      (Fr.  Berceau   du   Moteur).     The 

timbers  or  fuselage  members  upon  which  the  engine  is  fastened. 
ENGINE   SPIDER  or  BRACKET.      (Fr.  Arraignee   Support   dc 

Moteur.')    A  perforated  metal  support  for  a  rotary  cylinder  motor. 
ENTERING  EDGE.     (Fr.  Bord  D'Attaque  or  Arctier  Avant.)    The 

front  edge,  or  air  engaging  edge,  of  an  aerofoil  or  lifting  surface. 

It  is  also  called  the  "Leading  Edge." 
ENTOILAGE  (Fr.)     Wing  fabric  or  covering. 
ENVELOPE.    The  gas  bag  of  a  balloon  or  dirigible. 
ESSIEU  (Fr.)     Axle. 
ENVERGURE   (Fr.)     Wing  span. 
EXPANDING  PITCH.     A  system  in  which  the  pitch  increases  or 

"expands"  towards  the  tips  of  the  propeller. 


FABRIC,  WING.     (Fr.  Entoilage.)     The  cloth  used  for  covering 

the  wing  and  control  surface  structures. 
FAIRING.     (Fr.  Fusele.)     Wood  coverings  used  to  streamline  steel 

struts  or  other  structural  members. 
FERRULES.     Sheet  steel  caps  used  for  the  ends  of  the  interplane 

struts. 
FIN.     (Fr.  Derive.)     A  fixed  vertical  stabilizing  surface  used  for 

damping  out  horizontal  vibration  and  oscillations. 
FINENESS  RATIO.     The  ratio  of  the  maximum  length  to  the 

width  of  a  streamline  body. 
FITTINGS.      (Fr.  Ferriires,  Godets.)     The  metal  parts  used   for 

making  connections  between  the  structural  parts  of  an  aeroplane. 
FIXED  TAIL.     (See  STABILIZER.) 
FLACCID   BLADE   PROPELLER.     A   propeller   having  a   cloth 

covered  frame  work  on  which  the  fabric  is  free  to  adjust  itself 

to  the  air  pressure. 
FLAPS,  ELEVATOR.     (Fr.   Volets  de  Profondeur.)     See  ELE- 
VATOR. 
FLEXIBLE    SHAFT.      (Fr.    Transmission    flexible.)      Used    for 

tachometer  drive. 
FLOORING.     (Fr.  Plancher.) 
FLASQUE  D'HELICE  (Fr.)     Propeller  flange. 
FOOT  LEVER.     (Fr.  Palomiier.)     The  foot  lever  generally  used 

to  operate  the  rudder. 


GLOSSARY 

FORMERS.  Supports  used  in  giving  a  certain  outline  to  the  fusel- 
age. The  formers  are  attached  to  the  fuselage  frame  and  in  turn 
support  small  stringers  on  which  the  fabric  is  fastened. 

FRICTIOXAL  WAKE.  The  following  current  of  air  in  the  rear 
of  a  moving  body  or  surface.  Because  of  the  friction,  a  portion 
of  the  air  is  drawn  in  the  direction  of  the  motion. 

FUSELAGE.  A  structure,  usually  enclosed,  which  contains  and 
streamlines  the  power  plant,  passengers,  fuel,  etc.  Sometimes 
called  the  "Body." 

FUSELAGE  BIPLANE.       See  TRACTOR  BIPLANE. 

FUSIFORM.     Of   streamline    form. 

G 

GAP.  The  vertical  distance  between  leading  edges  of  the  super- 
posed planes  of  the  biplane  or  triplane. 

GLIDING.  (Fr.  Vol  Plan.)  With  an  aeroplane  the  weight  of  the 
machine  can  be  made  to  provide  a  forward  component  that  will 
allow  the  machine  to  descend  slowly  (without  power)  along  an 
inclined  line.     This  line  is  known  as  the  "Gliding  Path." 

GLIDING  ANGLE.  The  angle  made  by  the  gliding  path  with  the 
horizontal  is  known  as  the  gliding  angle.  This  may  be  expressed 
in  degrees  or  in  the  units  of  horizontal  distance  traveled  per  foot 
of  fall. 

GLIDER.  A  small  form  of  aeroplane  without  a  power  plant,  which 
is  capable  of  gliding  down  from  an  elevation  in  the  manner  of 
an  aeroplane.  With  a  proper  direction  and  velocity  of  wind  it 
can  be  made  to  hold  a  constant  altitude  and  can  be  made  to  hover 
over  one  spot  continuously. 

GOUVERXAIL  (Fr.)     Rudder. 

GUY  WIRE.    A  bracing  wire. 

H 

HARD  WIRE.    A  solid  tempered  wire  of  high  tensile  strength  used 

for  aeroplane  bracing  systems. 
H.  T.  WIRE.    Another  expression  for  hard  or  high  tensile  strength 

wire. 
HEAD  RESISTANCE.     The  resistance  of  the  structural  parts  of 

an  aircraft.    In  an  aeroplane,  the  head  resistance  is  the  sum  of  the 

resistances  of  the  body,  stays,  struts,  chassis,  tail,  rudder,  elevators, 

etc. ;   in   fact,  this  includes  everything  with  the  exception  of  the 

wing  drag. 
HELICOPTER.     A  type  of  direct  lift  machine  in  which  sustenta- 

tion  is  performed  by  vertical  air  screws  or  propellers. 


GLOSSARY 


GLOSSARY 

HELIX.     A  geometrical  curve  formed  by  the  combined  advance  and 

revolution  of  a  point. 
HELICE  (Fr.)     Propeller  or  screw. 
HELICE  TRACTIVE  (Fr.)     Tractor  propeller. 
HYDROAEROPLANE.    See  SEAPLANE. 
HOLLOW  WOOD  CONSTRUCTION.     (Fr.  Bois  Creus.) 
HOOD  OF  ENGINE.     (Fr.  Capot.) 
HYDROMETER.     An   instrument    for    measuring   the    density   of 

liquids. 

I 

ICTHYOID.    Fish  or  stream  lined  shape. 

INCLINED  PLANE.    A  plane  inclined  to  the  wind  stream  so  that 

the  energy  of  the  air  stream  is  broken  up  into  the  two  components 

of  lift  and  drag. 
INCLINOMETER.    An  instrument  used  for  determining  the  angle 

of  the  flight  path. 
INCIDENCE.    See  ANGLE  OF  INCIDENCE. 
INCIDENCE,  NEGATIVE.    The  angle  formed  with  the  air  stream 

when   front  edge  of  the  lifting  surface  dips  below  the  apparent 

flight  path. 
INHERENT  STABILITY.     Stability  due  to  some  f\xed  arrange- 
ment of  the  main  or  auxiliary  surfaces.     A  machine  that  requires 

mechanism  or  moving  parts  for  its  stability  is  automatically  but 

not  inherently  stable. 
INTERFERENCE.     The  crowding  of  the  airstream  in  the  gap  of 

a  biplane  or  triplane  causes  the  surfaces  to  "Interfere,"  and  results 

in  a  loss  of  lift. 

J 

JOY  STICK.  See  CONTROL  STICK. 
JA}[B  de  FORCE  (Fr.)  Bracing  strut. 
J  ANTE  (Fr.)     Rim  of  wheel. 


KITE  BALLOON.     See  BALLOON. 

KILOMETER.  French  metric  unit  of  distance.  One  kilometer 
equals  0.621  statute  mile  or  0.5396  nautical  mile. 

KILOGRAM.  Metric  unit  of  weight.  One  kilogram  equals  2.205 
Avoir,  pounds. 

KNOCKOUT  HUB.  An  aeroplane  chassis  wheel  hub  provided 
with  removable  bronze  bushings. 

KEEL  PLANE  AREA.*  The  total  effective  side  area  of  an  aero- 
plane which  tends  to  prevent  skidding  or  side  slipping. 


GLOSSARY 


LATERAL  STABILITY.     Stability  about  the  fore  and  aft  axis. 

L' AVI  ON  (Fr.)     Aeroplane. 

LAMINATED.    Built  up  in  a  series  of  layers. 

LEEWAY.*  The  angular  deviation  from  a  given  course  due  to  cross 
currents  of  wind. 

LEADING  EDGE.     See  ENTERING  EDGE. 

LIFT.  The  vertical  component  of  the  forces  produced  on  an  aero- 
foil by  an  air  current. 

LIFT  COEFFICIENT.  The  lift  per  unit  of  area  at  a  unit  velocity 
(Ky).  The  American  lift  coefficient  is  the  lift  in  pounds  per  square 
foot  at  one  mile  per  hour. 

LIFT  CAPACITY.    See  CAPACITY. 

LIVE  LOAD.  The  live  load  generally  includes  the  passengers, 
pilot,  fuel,  oil,  instruments,  and  portable  baggage,  although  in 
some  cases  the  instruments  are  included  in  the  dead  load.  The 
live  load  is  the  difference  between  the  total  lift  and  the  dead  load. 

LOADING  (UNIT).  The  unit  loading  is  the  load  carried  per 
square  foot  of  wing  surface,  or  is  equal  to  the  total  weight  divided 
by  the  area. 

LONGERONS.  The  principal  fore  and  aft  structural  members  of 
the  fuselage. 

LONGITUDINALS.    See  LONGERONS. 

LONGITUDINAL  STABILITY.  Stability  in  a  fore  and  aft  direc- 
tion about  the  "Y"  axis. 

M 

MASS.  The  quantity  of  matter.  Is  equal  to  the  weight  in  pounds 
divided  by  the  gravitation,  or  generally  to  the  weight  divided  by 
32.16. 

MANDRIN  de  BOIS  (Fr.)    Wood  spar. 

MAROUFLAGE   (Fr.)     Strut  taping  with  fabric  bands. 

MAN  CHE  A  BALAI  (Fr.)     Control  stick. 

MANETTE  (Fr.)    Throttle. 

MAR,  MONTANT  (Fr.)     Interplane  struts. 

METACENTER.*  The  point  of  intersection  of  a  straight  vertical 
line  passing  through  the  center  of  gravity  of  the  displaced  fluid 
or  gas,  and  the  line  that  formerly  was  a  vertical  through  the 
center  of  gravity  before  the  body  was  tipped  from  its  position  of 
equilibrium.  There  is  a  different  metacenter  for  each  position  of 
a  floating  body. 

MONOPLANE.  (Fr.  Monoplan.)  A  type  of  aeroplane  with  a 
single  wing  surface. 


GLOSSARY 

MONOCOQUE  BODY  (Fr.)  A  body  built  up  in  tubular  form 
out  of  three-ply  wood,  thus  virtually  forming  a  single  piece  body. 

MONOPLACE  (Fr.)     Single  seater. 

MOXOSOUPAPE  (Fr.)     Single  valve  Gnome  motor. 

MONTGOLFIER   (Fr.)     Hot  air  balloon. 

MULTIPLANE.  An  aeroplane  having  the  main  lifting  surface 
divided  into  a  number  of  parts. 

N 

NACELLE.  The  body  or  fuselage  of  an  aeroplane  or  dirigible.  It 
generally  signifies  a  dirigible  body.  The  short  fuselage  of  a 
pusher  type  is  often  called  the  nacelle. 

NATURAL  STABILITY.    See  INHERENT  STABILITY. 

NEGATIVE  AILERONS.  Ailerons  making  a  negative  angle  with 
the  wind  when  in  normal  flight.  The  negative  incidence  of  the 
ailerons  is  decreased  on  the  low  side  and  increased  on  the  high 
side  so  that  the  high  side  is  pushed  down.  This  decreases  the 
drag  on  the  lower,  inner  wing  in  making  a  turn,  and  therefore 
does  not  tend  to  stall  the  machine. 

NERVURES  (Fr.)     Wing  ribs. 

NORMAL  PLANE.  A  flat  plane  placed  with  its  surface  at  right 
angles  to  the  air  stream. 

NORMAL  PRESSURE.  The  pressure  at  right  angles  to  the  sur- 
face of  a  plane. 

NON-LIFTING  TAIL.  A  stabilizing  surface  arranged  so  that  it 
carries  no  load  in  normal  flight. 

NOSING.    The  member  used  for  the  entering  edge  of  the  wing. 

NOSE.    The  front  end  of  the  aeroplane. 

NUT.     (Fr.  Ecru.) 

O 

OBLITEUR  RINGS  (Fr.)  The  special  piston  rings  used  on  the 
Gnome  motor. 

ORTHOPTER.     Any  type  of  wing  flapping  machine. 

ORNITHOPTER.  A  wing  flapping  machine  that  imitates  bird 
flight. 

ORTHOGONAL  BIPLANE.  A  biplane  with  the  upper  and  lower 
leading  edges  in  line. 

OSMOSIS.  The  transfer  of  hydrogen  or  other  gas  through  a  bal- 
loon envelope  by  a  molecular  process.  This  must  not  be  confused 
with  leakage  due  to  holes. 

OUTRIGGER.    See  Boom. 


GLOSSARY 

O.  W.  L,  TYPE.    A  type  of  machine  adapted  for  use  over  "Water 
and  Land." 


PANCAKE.    A  straight  vertical  drop  due  to  stalling. 

PATH  OF  FLIGHT.  The  path  of  the  center  of  gravity  of  an  air- 
craft in  reference  to  the  air. 

PALONNIER  (Fr.)     Foot  bar  or  lever. 

PANELS.     The  wing  sections  included  between  adjacent  struts. 

PATIN,  PATINNAGE  (Fr.)     Skids. 

PARASOL  TYPE.  A  monoplane  in  which  the  wing  is  located 
above  the  body. 

PENGUIN.    A  training  machine  which  cannot  leave  the  ground. 

PERSONNEL.    Pilot  and  passengers. 

PETROL.    An  English  term  for  gasoline. 

PHILLIP'S  ENTRY.     See  DIPPING  FRONT  EDGE. 

PILOT.    The  operator  of  aircraft. 

PILOT  BALLOONS.  Small  balloons  sent  up  to  determine  the 
direction  of  the  wind. 

PITCH.  The  forward  distance  traveled  through  by  one  revolution 
of  the  propeller. 

PITCHING.  A  fore  and  aft  oscillation,  first  heading  up  and  then 
diving. 

PISCIFORM.    Fish  form. 

PIQUE,  VOL  (Fr.)     Dive. 

PLAFOND   (Fr.)     "Ceiling"  or  maximum  altitude  obtainable. 

PLAN  CENTRAL  (Fr.)     Center  panel. 

PLANCHER  (Fr.)     Flooring. 

PLAN  FIXE  de  QUEUE  (Fr.)     Stabilizer  surface. 

PLAN  de  DERIVE  (Fr.)     Stabilizing  fin. 

PITOT  TUBE.  An  instrument  for  measuring  the  velocity  of  an 
air  current. 

PONTOON.     Seaplane  floats. 

POIGNEE  (Fr.)     Handle. 

POMPE  (Fr.)     Pump. 

POMPE  A  PRESS  I  ON  (Fr.)     Pressure  pump. 

POULIE    (Fr.)      Pulley. 

PROPELLER.  (Fr.  Hclice.)  A  device  used  in  converting  the 
energy  of  a  motor  into  the  energy  required  for  the  propulsion  of 
an  aircraft.  It  consists  of  two  or  more  rotating  blades  which  are 
inclined  in  regard  to  the  relative  wind,  and  hence  they  act  as 
rotary  aeroplanes  in  creating  a  tractive  force. 


GLOSSARY 

PROPELLER  ROTATION.  The  direction  of  rotation  is  deter- 
mined when  standing  in  the  slip  stream. 

PNEU   (Fr.)     Pneumatic  tire. 

PTERYGOID  ASPECT.  A  wing  flying  with  the  long  edge  to  the 
wind  is  said  to  be  in  "Pterygoid  Aspect." 

PUSHER  TYPE.  An  aeroplane  with  the  propeller  in  the  rear  of 
the  wings. 

PYLON.    A  marking  post  on  an  aeroplane  course. 


RACE    OF   A    PROPELLER.      The    air    stream    thrown    by    the 

propeller. 
RADIAL  MOTOR.    A  motor  with  the  cylinders  arranged  in  radial 

lines  around  the  crankcase. 
RAKED  TIPS.     The  tips  are  arranged  at  an  angle  with  the  wing 

so  that  the  span  of  the  trailing  edge  is  greater  than  that  of  the 

leading  edge. 
RAYONS  (Fr.)     Spokes. 
REFLEX  CURVE.    An  aerofoil  in  which  the  trailing  edge  is  given 

an  upward  turn. 
REMOUS  (Fr.)     A  downward  current  of  air. 
RESERVOIR   (Fr.)     Tank. 

RESERVOIR  SOUS  PRESSION  (Ft.)     Pressure  tank. 
RESULTANT.     The  total  force  resulting  from  the  application  of 

a  number  of  forces. 
RETREAT.    Back  swept  wings  with  the  tips  to  the  rear  of  the  wing 

center. 
RIBS.     The  fabric  forming  member  of  the  wing  structure. 
RUDDER,  VERTICAL.     A  control  surface  used  for  steering  in  a 

horizontal  plane. 
ROLL.     Oscillation  about  the  fore  and  aft  axis. 


SCREW.     (Fr.  Helice.)     See  PROPELLER. 

SEAPLANE.*    An  aeroplane  equipped  with  floats  or  pontoons  for 

landing  on  water. 
SCOUT  TYPE.    See  CHASER. 
SERVICE  TANK.     The  fuel  tank  feeding  directly  into  the  carbu.- 

retors. 
SET  BACK  WINGS.     A  type  of  wing  in  which  the  leading  edge 

is  inclined  backward  as  in  the  Mann  biplane.    The  trailing  edge  is 

straight. 


GLOSSARY 

SHOCK  ABSORBERS.  An  elastic  device  on  the  chassis  or  landing 
gear  that  absorbs  vibration  by  allowing  a  limited  axle  movement. 

SIDE  SLIP.*  Sliding  down  sideways,  and  toward  the  center  of  a 
turn.    This  is  due  to  an  excessive  angle  of  bank 

SIEGE  (Fr.)     Seat. 

SIMILITUDE,  LAWS  OF.  The  drag  or  resistance  of  a  small 
aerodynamic  body  does  not  increase  in  direct  proportion  with  the 
area  and  speed.  The  laws  governing  the  relation  between  a  model 
and  a  full  size  machine  are  known  as  the  laws  of  "Similitude." 

SKIDS.*  (Fr.  Patin,  Pattinage.)  Long  wood  or  metal  runners  at- 
tached to  the  chassis  to  prevent  the  "nosing  over"  of  a  machine 
when  landing,  or  to  prevent  it  from  dropping  into  holes  or  ditches 
on  rough  ground.     It  also  acts  when  the  wheels  collapse. 

SKID  CURTAINS.  Vertical  side  curtains  or  surfaces  provided  to 
reduce  the  skidding  action  on  turns  or  to  prevent  side  slip. 

SKIDDING.*  Sliding  sideways  away  from  the  center  of  the  turn. 
It  is  due  to  insufficient  banking  on  a  turn. 

SKIN  FRICTION.  The  resistance  caused  by  the  friction  of  the  air 
along  a  surface. 

SLIP.*  Applied  to  propeller  action,  the  slip  is  the  difference  between 
the  actual  advance  of  an  aircraft  and  the  theoretical  advance  cal- 
culated from  the  product  of  the  mean  pitch  and  the  revolutions  per 
minute.  When  the  propeller  is  held  stationary,  the  slip  is  said  to 
be  100  percent. 

SLIP  STREAM.    The  wind  stream  thrown  by  a  propeller. 

SOARING  FLIGHT.*  The  sustentation  of  a  wing  surface  due  to 
wind  currents  and  without  the  expenditure  of  other  power.  Soar- 
ing flight  is  performed  by  gulls,  buzzards  and  vultures,  but  no 
practical  machine  has  yet  been  built  that  will  fly  continuously  with- 
out the  aid  of  power. 

SPAN.  (Fr.  Envergure.)  The  length  or  longest  dimension  of  a 
wing,  generally  taken  at  right  angles  to  the  wind  stream. 

SPAR.  (Fr.  Bras  D'Aile.)  The  main  wing  beams  that  transmit 
the  lift  to  the  body. 

SPOKES.     (Fr.  Rayon.) 

SPREAD.*     See  SPAN. 

STABILITY.*  The  property  of  an  aircraft  that  causes  it  to  return 
to  a  condition  of  equilibrium  after  meeting  with  a  disturbance  in 
flight. 

STAGGER.*  The  advance  of  the  leading  edge  of  the  upper  wing 
over  that  of  the  lower  wing. 


GLOSSARY 

STABILIZER.*  (Fr.  Stabilisateur.)  A  horizontal  tail  surface 
(fixed)  used  for  damping  out  oscillations  and  for  promoting  longi- 
tudinal stability. 

STALLING.*  The  condition  of  an  aeroplane  that  has  lost  the 
speed  necessary  for  steerage  way  or  control. 

STALLING  ANGLE.     See  CRITICAL  ANGLE. 

STEERING  WHEEL.     (Fr.  Volant.) 

STATOSCOPE.*  An  instrument  for  detecting  a  small  rate  of  as- 
cent or  descent.     Used  principally  with  balloons. 

STAY  WIRE.  (Fr.  Tendeiir.)  A  wire  or  cable  used  as  a  tie  to 
hold  members  together,  or  to  give  stiffness  to  a  structure. 

STEP.*     A  break  in  the  form  of  a  float  or  flying  boat  bottom. 

STREAMLINE.  A  form  of  body  that  sets  up  no  turbulence  or 
eddies  in  passing  through  air  or  liquid. 

STRUT.*  (Fr.  Mar,  Montant.)  A  compression  member  used  in 
separating  the  upper  and  lower  wings  of  a  biplane,  or  the  longerons 
of  the  fuselage. 

SWEEP  BACK.     See  RETREAT. 


TACHOMETER.  (Fr.  Comptc  Tours.)  An  instrument  for  di- 
rectly indicating  the  revolutions  per  minute. 

TAIL.*  (Fr.  Queue.)  The  rear  part  of  an  aircraft  to  which  usu- 
ally are  attached  the  rudder,  stabilizer,  and  elevator. 

TAIL  SKID.  (Fr.  Bequille.)  A  flexibly  attached  rod  which  holds 
the  tail  surfaces  off  the  ground,  and  breaks  the  landing  shocks  on 
the  tail  structure. 

TAIL  BOOAI.     See  BOOM. 

TAIL  DIVE.    A  very  dangerous  backward  dive. 

TAIL  SPIN.  A  condition  in  which  the  tail  revolves  about  a  vertical 
line  passing  through  the  center  of  gravity. 

TANDEM  PLANES.  A  form  of  aeroplane  in  which  the  wings  are 
placed  one  after  another. 

TAUBE.  An  old  type  of  German  or  Austrian  aeroplane  with  back 
swept  wing  tips. 

TAXI.     To  run  along  the  ground. 

THIMBLE.  (Fr.  Cosse.)  An  oval  grooved  metal  fitting  used  for 
the  protection  of  a  cable  loop  at  the  point  of  attachment. 

THREE-PLY.  (Fr.  Contre plaque.)  A  wood  sheet  composed  of 
three  layers  of  wood  glued  together,  the  line  of  grain  crossing  at 
each  joint. 

THRUST.    The  propulsive  force  exerted  by  a  propeller. 


GLOSSARY 

THRUST  DEDUCTION.*  The  reduction  of  thrust  due  to  a  reduc- 
tion of  pressure  under  the  stern  of  the  aircraft. 

TIRANT  (Fr.)     Bracing  tubes. 

TORQUE.     The  turning  force  or  moment  exerted  by  the  motor. 

TOILE   (Fr.)     Linen. 

TORQUE  WARP.  The  amount  of  warp,  or  permanent  set  in  the 
ailerons  necessary  to  overcome  the  torque  or  twisting  effect  of  the 
motor.  In  some  machines  the  torque  is  overcome  by  changing  the 
angle  of  incidence  at  the  wing  tips. 

TRACTOR  BIPLANE.  A  type  of  aeroplane  in  which  the  propeller 
is  placed  in  front  of  the  wings  so  that  it  pulls  the  machine  along. 

TRAILING  EDGE.  The  edge  of  a  wing  at  which  the  air  stream 
leaves  the  surface. 

TRAIN  D'ATTERRISSAGE  (Fr.)     Landing  gear. 

TRANSFIL  (Fr.)     Cord  winding  on  the  struts. 

TRIPLANE.  (Fr.  Triplan.)  An  aeroplane  with  three  superposed 
wings. 

TURBULENCE.  The  eddies  or  discontinuity  caused  by  a  body  or 
surface  passing  through  the  air. 

U 
USEFUL  WEIGHT.     The  difference  between  the  total  lift  and  the 
dead  load.     This  comprises  the  pilot  and  passenger,  the  weight  of 
the  fuel,  baggage  and  instruments. 
UNIT   LOADING.    The   weight   per   square    foot   of   main   wing 
surface. 

V 

VOL  PLANE  (Fr.)     See  GLIDE. 

VOL  PIQUE  (Fr.)     See  DIVE. 

VOLANT  (Fr.)     Steering  wheel. 

VERNIS  (Fr.)     Varnish. 

VRIL  (Fr.)     Spinning  nose  dive. 

VOLETS  de  PROFONDEUR  (Fr.)     Elevator  flaps. 

W 
WARP  CONTROL*    Lateral  control  obtained  by  twisting  the  wing 

tips. 
WASHOUT.     Decreased  camber  or  incidence  toward  the  wing  tips. 
WEATHER-COCK  STABILITY.     Stability  in  the  line  of  travel,  or 

with  the  relative  wind,  so  that  the  machine  always  tends  to  head 

into  the  wind. 
WHALE  TYPE.     A  speed  type  biplane  in  which  the  body  entirely 

fills  the  gap  between  the  upper  and  lower  wings. 


GLOSSARY 

WHIRLING  TABLE.  A  testing  device  in  which  a  model  wing  or 
body  is  placed  at  the  end  of  a  revolving  arm. 

WINGS.*  (Fr.  Ailc.)  The  main  supporting  surfaces  of  an  aero- 
plane. 

WORKING  FACE.  The  face  of  a  propeller  blade  lying  next  to  the 
slip  stream. 

WAKE  GAIN.*  Due  to  skin  friction  and  eddies,  a  moving  aircraft 
drags  a  certain  amount  of  the  surrounding  air  with  it.  This  re- 
duces the  effective  resistance  of  the  hull  and  increases  the  effective 
pitch  of  a  pusher  propeller  since  the  latter  acts  on  a  forward 
moving  mass  of  air.     This  is  "Wake  gain." 


Launching  with  Catapult   from  Deck   of  Battleship. 


GLOSSARY 


GLOSSARY 


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IP 8  Wire 
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Pi/t/Crtess  of  Thimt/c 


Standard  U.   S.  A.  Wire  and  Cable  Connections.     Top  Shows  Stranded  H.  T. 

Cable  with  Thimble  in  Loop  and  Soft  Wire  Serving  Bands.      Center  Is  Solid 

Hard  Wire  Loop.     Bottom  Is  Detail  of  Thimble. 


GLOSSARY 


Progressive    Steps    in    Applying    Sewing    to    Stranded    Cable    (No    Thimble). 

After  Sewing   Is    Applied    the    Connection    Is    Soldered    All    Over   As    Shown 

by    (d).      Same   Applies   When   Thimble   Is    Used. 


(a)    and    (b)    Show   Hard    Wire    Connections   Made   with   Copper   Sleeve,      (c) 

and   (d)  Show  "Turk's  Head"  Connection  of  Stranded  Cable,   in  Which  End 

of  Cable   Is   Flayed  Out   and   Soldered  into  Connection  Socket. 


h  .  3-2-^ 


University  of  California 

SOUTHERN  REGIONAL  LIBRARY  FACILITY 

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